PRELIMINARY DESIGN REVIEW November 4, 2016

Transcription

1 NASA Student Launch Competition PRELIMINARY DESIGN REVIEW November 4, 206 California State Polytechnic University, Pomona 380 W Temple Ave, Pomona, CA 9768 Department of Aerospace Engineering Proposal September 30, 206

2 PURPOSE The Preliminary Design Review demonstrates that the overall preliminary design meets all requirements with acceptable risk, and within the cost and schedule constraints, and establishes the basis for proceeding with detailed design. It shows that the correct design options have been selected, interfaces have been identified, and verification methods have been described. Full baseline cost and schedules, as well as all risk assessment, management systems, and metrics, are presented Student Launch Handbook

3 Table of Contents Acronyms General Information Advisors and Mentors Team Members National Association of Rocketry Website Compliance Summary of Project Team Summary Launch Vehicle Summary Size and Total Mass Motor Choice Recovery System Subscale Launch Vehicle Summary Payload Experiment Summary Milestone Review Flysheet Proposal Changes Vehicle Changes Payload Changes Project Plan Changes Launch Vehicle Criteria Mission Statement and Mission Success Criteria Preliminary Vehicle System Design Preliminary Vehicle System Trade Study Leading Preliminary Vehicle System Design Launch Vehicle Subsystem Overview Launch Vehicle Subsystem Trade Studies Propulsion Subsystem Aerodynamics Subsystem

4 4.3.3 Avionics Subsystem Launch Vehicle Leading Design and Mass Summary Recovery Subsystem Recovery Subsystem Overview Main Parachute Recovery Component Trade Study Drogue Parachute Recovery Component Trade Study Recovery Avionics Trade Study Recovery Subsystem Redundancy Parachute Sizing and Safe Decent Analysis Recovery Subsystem Leading Components Deployment Charge and Altimeter Layout Recovery Bay Overview Full Scale Launch Vehicle Interfaces and Integrations Interface A: Avionics Integration Interface B: Fin Integration Interface C: Launch Vehicle and Ground Station Mission Performance Predictions Mission Performance Criteria Flight Profile Simulations Altitude Predictions Motor Thrust Curve Launch Vehicle Load Verification Stability Margin Calculated CP and CG Kinetic Energy Drift Calculations (no wind, 5 mph, 0, 5, 20) Preliminary Launch Plan Overview Test Launch, Pre-Launch and Launch Compliance Checklist Preliminary Checklist of Final Assembly and Launch Procedures Recovery Preparation

5 5.2.2 Scientific Payload Preparation Motor Preparation Final Assembly and Launch Preparations Payload Criteria Primary Payload Experiment Payload Objective and Experiment Success Criteria System Trade Study System Leading Design Leading Mass Summary Leading Drawings and Electrical Schematics Interfaces: Payload and Launch Vehicle Precision, Repeatability, Recovery Secondary Payload Experiment Payload Objective and Experiment Success Criteria System Trade Study System Leading Design FMP Mass Summary Drawings and Schematics Interfaces: Payload and Launch Vehicle Precision, Repeatability, Recovery Payload Standard Preliminary Operating Procedure Safety Safety Officer Identified Responsibilities Risk and Delay Impact Failure Modes and Effects Analysis (FMEA) Personnel Hazard Analysis Hazardous Materials Handling and Operations Hazard Recognition and Accident Avoidance

6 7.4.3 Carbon Fiber Composites Other Construction Materials Pre-launch Briefing Launches Unmanned Rocket Launches and Motor Handling Compliance with Federal, State, and Local Laws Written Compliance to Safety Regulations Caution Statement Plans, Procedures, and Other Working Documents Machining Lab Environmental Concerns and Effects Project Plan Requirements Compliance Matrix Launch Vehicle Compliance Matrix Recovery System Compliance Matrix Payload Compliance Matrix Safety Compliance Matrix General Compliance Matrix Derived Requirements Compliance Matrix Preliminary Budget Plan Full Scale Launch Vehicle Expenses Sub Scale Launch Vehicle Expenses Recovery System Expenses Payload Expenses Educational Engagement Expenses and Other Travel Expenses Overall Expenses Funding Plan Preliminary Timeline Timeline Gantt Chart

7 9.0 Educational Engagement Plan Educational Engagement Status Educational Engagement Summary and Details

8 Acronyms AGL = Above Ground Level AIAA = American Institute of Aeronautics and Astronautics CDR = Critical Design Review CDR = Critical Design Review CFR = Code of Federal Regulations CG = Center of Gravity CPP = California State Polytechnic University, Pomona Dps = degree per second FAA = Federal Aviation Administration FLVC = Full Scale Launch Vehicle Completion FMP = Fragile Material Protection Payload FN = Foreign National FRR = Flight Readiness Review g = acceleration due to gravity GPS = Global Positioning System GUI = Graphical User Interface IIS = Ignition Insertion System IMU = Inertial Measurement Unit KSI = kilo pound per square inch LLC = Limited Liability Company LRR = Launch Readiness Review LSB = Least Significant Bit LV = Launch Vehicle MATLAB = Matrix Laboratory MSDS = Material Safety Data Sheet NAR = National Association of Rocketry NASA = National Aeronautics and Space Administration NFPA = National Fire Protection Association NSL = NASA Student Launch PC = Payload Completion PDR = Preliminary Design Review PPE = Personal Protective Equipment PRA = Payload Retrieval Arm PRE = Payload Retrieval Elevator PRS = Payload Retrieval System RAC = Risk Assessment Codes 7

9 RAL = Rocket Assembly Laboratory RCF = Refractory Ceramic Fiber RCS = Reactionary Control System Payload RIS = Roll Induction System RSO = Range Safety Officer SCRA = Southern California Rocket Association SL = Student Launch SLVC = Subscale Launch Vehicle Completion STEM = Science, Technology, Engineering, and Mathematics TRA = Tripoli Rocketry Association UMBRA = Undergraduate Missiles and Ballistics Rocketry Association WBS = Work Breakdown Structure WE = Website Established 8

10 .0 General Information. Advisors and Mentors Donald L. Edberg, Ph.D Dr. Donald Edberg will be the team s faculty advisor. With about thirty years of experience combined in the aerospace industry and teaching, he will provide great advice on the project overall. Dr. Edberg has been employed at General Dynamics, AeroVironment, McDonnell Douglas, and the Jet Propulsion Laboratory. He was also formerly a Boeing Technical Fellow at both Boeing s Phantom Works and Boeing Information, Space and Defense Systems and most recently a faculty fellow at NASA Marshall Space Flight Center. Dr. Edberg has advised numerous senior projects at his time in California Polytechnic University, Pomona for the aerospace engineering department. In addition, he has been an advisor for the past four years for the NASA Student Launch Initiative teams. Rick Maschek rick.maschek@rocketmail.com Rick Maschek will act as one of the mentors for the team. Rick is heavily experienced in rocketry with over fifty years of flying rockets and he has built his own hybrid rockets. He holds a Tripoli Rocketry Association Level 2 certification and his certification number is #388. Since he holds a certification, he will perform all handling of igniters, motors, and charges. Todd Coburn, Ph.D. tdcoburn@cpp.edu Dr. Todd Coburn will act as the other mentor for the team. Dr. Coburn has been employed in Us Technical Consultants, Boeing Commercial Airplanes, and at Boeing Defense Systems; totaling to nearly thirty years of experience in the aerospace industry, specifically with strength analysis. Dr. Coburn will provide much insight on the launch vehicle s structural properties and capabilities. 9

11 .2 Team Members Figure.2- Team Members WBS 0

12 Figure.2-2 Task Forces WBS Figure.2- shows the Work Breakdown Structure for the entire NASA Student Launch Team, CPP NSL. Team Lead, Samuel Daugherty, is shown at the top of the hierarchy, right above the Safety Officer and Deputy. There are three specific functional sub-teams: Structures, Aerodynamics, and Avionics; aiding the overall core team is the support team. In addition to each members sub-team responsibility, they also have a main Task Force that require participation in and attention. These tasks include: Outreach, Safety, Primary Payload, Secondary Payload, Recovery, and Website Managing. Figure.2-2 outlines the Task Forces. Each team member is given a responsibility and therefore will be active and engaged during the entirety of the project. Each team member is also required to participate in the design, construction, writing processes and as well as participation during presentations and flight preparations. However, the handling of the black powder and anything of that variant will be handled by the team s mentor. In addition, Jairo Sanchez is our sole foreign national team member.

13 .3 National Association of Rocketry The designated NAR/TRA section that the team will be working with for launch assistance is the Southern Rocket Association (SCRA) #430. The main launch site for the team that is covered under section #430 is at Lucerne Dry Lakebed. Martin Bowitz, SCRA president, is the main contact for this section; below is his contact information. Martin Bowitz Section #430 PO Box 565 Fullerton, CA Website Compliance The team website is located at as shown in Figure.4-. In addition to the primary function of hosting the team s deliverables in PDF format, it will keep the public up to date on the project, inform educators of outreach opportunities, host pictures and videos during the design and manufacturing process, link to social media outlets, host team member pictures and bios, and contain documentation from Cal Poly Pomona teams of past years. The website also features information and pictures on the rocketry club on campus the team works closely with, UMBRA. To create the Cal Poly Pomona NASA Student Launch Website, the Weebly website builder was used. Weebly allows the creator to build a website using a drag and drop interface editor. This ensures the ability to create a clean and professional website. In addition, Weebly comes with a free hosting platform. This cloud-based infrastructure provides fluid interaction for visitors. Figure.4-: home page 2

14 2.0 Summary of Project 2. Team Summary The California Polytechnic State University, Pomona NSL Team is a student organization located in Pomona, California. The project faculty advisor is Dr. Donald Edberg, Professor of Aerospace Engineering. The team s Rocketry mentor is Rick Mascheck (Tripoli Rocketry Association Level 2: #388). The team s structural engineering advisor is Dr. Todd Coburn, Professor of Aerospace Engineering. The mailing address of the department is as follows: California State Polytechnic University, Pomona Department of Aerospace Engineering 380 W Temple Ave. Pomona, CA Launch Vehicle Summary A graphic summary of the launch vehicle can be found below in Figure Figure 2.2-: Launch Vehicle Summary 2.2. Size and Total Mass The length of the launch vehicle from the tip of the nose cone to the end of the motor bay is 7.3 ft. The outer diameter of the body tube is 6.08 in and the total mass of the launch vehicle including a loaded motor is 28. lbs Motor Choice The launch vehicle utilizes an Aerotech L-50P motor, whose documented performance allows for reasonable estimates for the apogee, launch acceleration, and launch rail exit velocity Recovery System The recovery system, utilizing altimeters, activate at apogee by firing the aft ejection charges to release the cruciform drogue parachute. Once the launch vehicle has been decelerated, stabilized, and reached an altitude of 500 ft. the fore ejection charges will fire, releasing the toroidal main parachute Subscale Launch Vehicle Summary The subscale launch vehicle will be designed to be a one-half scale as compared to the full-scale launch vehicle. Its outer diameter will be 3 in. and the overall length will be approximately 3.52 ft. Additionally the subscale motor will be selected such that the thrust to weight ratio of both 3

15 the full-scale and subscale launch vehicles is roughly equivalent. This ensures that the subscale can appropriately reflect the loads the full-scale would experience during Ascent. 2.3 Payload Experiment Summary The primary payload takes the form of a Roll Induction System (RIS). Specifically, the RIS primarily consists of an autonomous aileron system which, following motor burnout, initiates two complete rotations and a counter rotation that ceases all angular displacement instigated by the active system. The secondary payload consists of a pill housing suspended within a payload bay whose purpose is to shield the provided fragile materials from the loads and impulses generated by the launch and recovery of the launch vehicle. 2.4 Milestone Review Flysheet Can be found at 4

16 3.0 Proposal Changes 3. Vehicle Changes Since PDR, it was decided that the tapered section of the body tube would be removed completely. The purpose of the tapered section of the body tube was to decrease the mass of the motor bay and to decrease the drag of the rocket. The motor bay had only contained the motor and fin integration as of the proposal but since then, see section 3.2, the payload for the RIS has been changed from a flywheel design to an actuated fin design. Previously the extra space for the motor bay was un-needed therefore the tapered design was to remove the extra space. The new payload design requires more space in the motor bay for electronics and servos necessary for an actuating the fin, therefore the body tube of the rocket will stay a constant 6 in diameter. Figure 3.- shows that the transition piece has been taken out for the benefit of adding more room for the equipment. The previous rocket body tube diameter of 4 in has now been changed to a constant diameter body tube. Tapered Section Non-Tapered Section Figure 3.- Old Design (left) New design (Right) Another aspect of the rocket that has been changed is the fin design. In the proposal, the original fin design focused on a trapezoidal fin. The purpose of this design was to limit the area of the fin that is exposed to the ground during the soft landing. By reducing the amount of fin close to the end of the rocket, it reduces the amount of energy that will be transferred to the fins and thus might prevent a critical failure of the fin. Since the fins were stationary with no actuating surfaces, this was the leading design for the proposal. For the PDR, a new fin design and location has been used in which a clipped trapezoidal fin will be used. This clipped trapezoidal fin offers the greatest benefit to our RIS payload design with the controllable surface on the fins. The trapezoidal fin will still be used however it will be clipped in the back end to a 90-degree angle. The reasons for this will make manufacturing the actuating surface on the fin much easier. Allowing a straight actuating surface to design and manufacture will make the interface connections less complicated and more seamless. This design does have disadvantages in which the trailing edge of the fin now has the potential of breaking during landing. This will be alleviated by moving the fins 2 in. forward from the bottom of the rocket. This should allow more clearance 5

17 during landing, while also allowing the actuating surfaces on the fin to function properly. This fin design change can be seen in Figure Figure 3.-2 Old Trapezoidal Fin design (left), New Clipped Trapezoidal Fin Design (right) The original design for the nose cone was set as an elliptical nose cone. This design was used as a base design to start calculations, but simulations were not used to confirm that this was a more efficient design. After using simulation software ANSYS, it was found that a ogive nose cone was a more efficient design. We found a decrease in drag of about 5% while using the ogive design, which is a substantial amount. This design change can be seen in figure Figure 3.-3 Old Elliptical Nose Cone (left) New Ogive Nose Cone (right) 3.2 Payload Changes There have been two major changes to the rocket s payload: ) Preliminary Avionics Package has been split into two separate, independent systems: Data Collection System (DCS): Responsible for acquisition and transmission of flight data to ground station. Payload Control System (PCS): Responsible for actuation of primary payload system pursuant to Req ) The Primary Payload Experiment leading design has been shifted from a momentum wheel concept to a fin-based control surface system. 3.3 Project Plan Changes There were some changes after the proposal submission made by the NASA Officials: Teams awarded moved to October 4, Kickoff and PDR Q&A moved to October 9, Team website established on October 3, PDR submission moved to November 4, PDR Video Teleconference moved to November 7, and the CDR Q&A due December 2. In addition, to compensate for weather mishaps and further delays, the full scale launched are moved to earlier dates. These changes were noted on the team Gantt chart located in section

18 4.0 Launch Vehicle Criteria 4. Mission Statement and Mission Success Criteria Our mission is to successfully build, test, and fly, a launch vehicle carrying multiple scientific payloads, allowing us to collect valuable data in flight. The fist payload is roll induction system (RIS) that will enable the launch vehicle to actively initiate two complete rotations and a counter rotation to cease all effects of the active roll mechanism post motor burnout. The second payload is a fragile material protection system (FMP) that will shield an unspecified payload from the forces resulting from the launch and recovery of the launch vehicle. In addition to both payloads, an avionics package will be collecting a suite of kinetic and atmospheric data in flight. 7

19 4.2 Preliminary Vehicle System Design 4.2. Preliminary Vehicle System Trade Study Launch Vehicle System Trade Study Alternative # - Non-tapered Alternative #2 - Tapered Description Made with a mix of 6 Diameter Blue Tube 2.0/Carbon Fiber for body 3-D printed/carbon Fiber reinforced fins/nosecone Ogive Nosecone Clipped trapezoidal fins Total length approximately 80 PROS: CONS: Single diameter, easier manufacture More familiar design Plenty of room to work with Unused space near motor mount Additional drag with 6 diameter Geared towards RCI-B & C Payload Design 6 Body tube pretty standard Most cons mitigated by getting slightly larger motor Description Mix of Carbon Fiber/Blue Tube 2.0 Tapered end for motor mount, D printed/carbon Fiber reinforced fins/nosecone Ogive Nosecone Clipped trapezoidal fins Total length approximately 00 PROS: Reduced drag and mass Smaller motor easier to mount More space efficient Easy implementation of observation bay CONS: Creates stress point Unfamiliar design Geared towards RCI-A Payload Design Design reduces induced drag The transition point is not as risky structurally than it seems 8

20 Team is more familiar with a single diameter tube design Leading Preliminary Vehicle System Design Although not familiar the tools needed to manufacture are available Leading Alternative: Alternative # The non-tapered design is currently the favored choice for functionality and feasibility when considering our leading payload design. The transition piece adds unnecessary complexity and the greater volume of the motor bay allows for the placement of servos to actuate the ailerons. Taking out the transition piece also removes the transition point as a critical stress point. This design also allows more room for the servos to be used on the RAS-B system. Main Parachute Secondary Payload Primary Payload Recovery Avionics Drogue Parachute Observation Motor Figure Leading Alternative Layout 9

21 4.2.3 Launch Vehicle Subsystem Overview Figure shows the preliminary launch vehicle system, subsystems and components in a product WBS. The text highlighted in red need to be discussed in the trade studies. Components and subsystems are not final and will be revised further in the design. Figure Product System WBS with Subsystems and Components 20

22 4.3 Launch Vehicle Subsystem Trade Studies 4.3. Propulsion Subsystem Table Propulsion Subsystem Trade Study Alternative # Alternative #2 Alternative #3 Aerotech L50P RMS 75/3840 Propellant: APCP Burn Time: 3.7 s Mass: oz Length: 20.9 in LV Apogee: 5556 ft PROS: Achieves altitude Altitude margin Easily Reloadable CONS: Greatest cost This architecture is the most expensive, however, in the long run, this motor has greater compatibility and ease of use bring the other costs down. In addition, it has a comfortable cushion for the altitude. This allows room for error and additional adjustments further along the line. Gorilla L425WC RMS 75/3500 Propellant: APCP Burn Time: 8.00 s Mass: 24 oz Length: 9.6 in LV Apogee: 5502 ft PROS: Achieves altitude Altitude margin Shorter length Less strain on avionics CONS: Incompatible with commercial retainer Requires additional tool for snap rings This architecture is a capable motor. It has a shorter required length leaving more room for other components. In addition, its long burn time reduces potential damage and inaccuracy in the avionics. Nevertheless, this motor is not compatible with commercial retainers and would drive up required costs and time. Gorilla L060G RMS 76/3500 Propellant: APCP Burn Time: 3.47 s Mass: 24 oz Length: 9.6 in LV Apogee: 5672 ft PROS: Achieves altitude Altitude margin Shorter length CONS: Incompatible with commercial retainer Requires additional tool for snap rings This engine s performance almost mirrors those of the Aerotech motor. It is both cheaper and shorter, however, like the L425WC, it also suffers from compatibility issues driving up costs. 2

23 Leading Alternative: Architecture # The Aerotech L50P motor is the optimal propulsion system for the launch vehicle. Despite its high initial costs, the motor satisfies the altitude requirement with a sizable cushion and offers both compatibility and usability. It provides options to utilize commercial retainers and threaded closures decreasing additional tools and increasing efficiency. Overall, the motor does not have the best affordability, but it achieves our target altitude within a reasonable margin and is easily reloadable. Mass Summary Figure Aerotech L50P Table Aerotech L50P Propulsion Subsystem Mass Summary Part Mass (oz) Motor Retainers 3.88 Motor Mount 90 22

24 4.3.2 Aerodynamics Subsystem Nose Cone Table Nose Cone Component Trade Study Alternative # Alternative #2 Alternative #3 Ogive Nose Cone Diameter: 6 in Length: 2 in Fineness Ratio: 2 CD = Volume = 0.06 ft 3 Mass = oz PROS: Lowest drag coefficient More stability CONS: Heaviest mass More material required Longer manufacturing Weak at tip This architecture has the most mass and volume and the least. It also provides the least drag coefficient making the nose cone more efficient in the subsonic range. However, the nose cone tip is susceptible to damage and the production may be slightly longer. Power Series.75 Nose Cone Diameter: 6 in Length: 2 in Fineness Ratio: 2 CD = Volume = ft 3 Mass = oz PROS: Least stability Lightest mass Least materials needed CONS: Highest drag coefficient Weak at tip This architecture is the lightest of all the designs. With this weight, the nose cone is substantially smaller in volume giving this design the fastest production. Like architecture, the nose cone tip is sharp and fragile. Power Series.5 Nose Cone Diameter: 6 in Length: 2 in Fineness Ratio: 2 CD = Volume = ft 3 Mass = oz PROS: Moderate drag coefficient Moderate stability Stronger with blunt tip CONS: Moderate mass More material required Longer manufacturing This architecture is a negotiation of the best and worst nose cone designs. It is slightly lighter than the architecture 2 and better aerodynamically. However, this nose cone is significantly smaller in volume preventing any avionic additions. 23

25 Leading Alternative: Architecture The ogive nose cone is the most beneficial to the performance of the launch vehicle. With the heavy motor weighing down the aft end, the center of gravity is pushed aft. This could cause a deficiency in the launch vehicle s stability. With a heavier nose cone and at a greater distance from the aft end, the stability easily increases with a lower required weight. The ogive nose cone is also the best design due to its low drag coefficient of In addition, it has the largest volume at 0. cubic feet allowing for compartments to be implemented if necessary. The ogive nose cone is a versatile architecture that satisfies performance, weight, and storage necessities. Figure Ogive Nose Cone Mass Summary Table Nose Cone Component Mass Summary Part Mass (oz) Ogive Nose Cone 56 24

26 Fins Table Fin Component Trade Study Alternative # Alternative #2 Alternative #3 Using the NACA0006 airfoil an un-swept fin was created with a 7.87in chord and 5.90in span PROS: Cd= Larger surface area allows for more room to integrate larger control surfaces and necessary equipment Contains the highest amount of usable volume for more servos and wiring to run through it CONS: Has highest drag Pressure is distributed heavily at the front Using the NACA0006 airfoil an un-swept fin was created with a 7.87in chord, 3.94 in tip chord, and 5.90in span and PROS: Cd= Has a more even pressure distribution taking stress off the leading edge Contains the second highest amount of internal volume for servos CONS: Doesn t have the lowest Cd Pressure distribution is still concentrated at the leading edge and is not uniform Using the NACA0006 airfoil an unswept fin was created with a 7.87in root chord, 3.94in tip chord, and 5.90in span and PROS: Cd= Has the lowest drag of all designs Pressure distribution is much more even resulting in a more inform flow Contains least amount of usable internal volume CONS: Has the least amount of workable internal area 25

27 All shapes were analyzed using the same dimensions and conditions Its high cd makes it undesirable from a drag perspective All shapes were analyzed using the same dimensions and conditions It also has the second lowest cd Good balance between volume and drag All shapes were analyzed using the same dimensions and conditions Alternative three possess the lowest Cd, but has the least amount of usable volume Leading Alternative: Alternative #3 Alternative two possesses a good balance between cd and usable volume, which is important to provide space for the aileron actuating mechanisms. The decision to use control surfaces means that many of the servos and wires will be routed inside the fins. This creates an imperative where internal volume must be considered. Drag must also be considered and again alternative two has the best balance between the two factors. Even though this design does not have the most optimal drag coefficient, it allows for simpler integration of our aileron actuating mechanism within the volume of the fin. Figure : Drawing of Leading Design root chord cross section 26

28 Figure Planform view of the Leading Design 27

29 Mass Summary Mass properties of NACA_0006 Swept Back Material: ABS Plastic 28

30 4.3.3 Avionics Subsystem Data Control Component The Data Control System (DCS) makes up the data logging and transmitting portion of the proposal s Preliminary Avionics Package. The objectives of this self-contained system are: Acquisition of accelerometer and gyroscopic flight data On-board storage of flight data Wireless transmission of flight data Table DCS Trade Study - CPU Components Arduino MEGA 2650 Raspberry Pi 3 Model B Arduino Micro The MEGA is one of the more powerful boards Arduino offers, featuring an ATmega2560 chip with 256 kilobytes of flash memory and 56 pins for PWM and analog devices. It also features four separate UART Serial busses for multi-component communication. PROS: 52 I/O Pins, PWM and Analog 256 KB of flash memory for programming Form factor allows use of component shields Programming simplicity CONS: Large form factor Overkill number of pins Not camera integration friendly The Raspberry Pi 3 Model B features a.2ghz 64-bit quadcore ARM Cortex-A53 CPU and GB RAM. In terms of raw processing power and magnitude of applications, the Raspberry Pi 3 Model B far surpasses the competition. PROS:.2 GHz 64-bit Processor Supports multiple programming languages 40 dedicated GPIO pins Camera support CONS: Large form factor Complexity lends to less durability The Arduino Micro is a small form factor 8 bit AVR microcontroller, similar in functionality to all Arduino controllers. With 20 I/O pins, it is less capable than the Mega for multi-component integration, however for our application it is more than sufficient. PROS: Small form factor Can be directly soldered onto PCB Programming simplicity CONS: Limited I/O capability Single serial interface Not component shield capable 29

31 Leading Alternative: Arduino MEGA The Arduino MEGA best satisfies the objectives of the Data Collection System. Since our project requires both an on-board data logging component and wireless transmission capability, an integration friendly platform is required. The MEGA is XBee Shield friendly, and can be easily mounted to a bulkhead within the rocket. It is also a more familiar platform, offering both ease of programming and troubleshooting. Should a backup GPS component be included in the DCS, the code on the MEGA can easily be accommodated. The larger form factor of the MEGA will fit within the inner diameter of our 6 body tube. It will interface with the bulkhead by utilizing the six mounting holes, as Figure below illustrates. Mass Summary of DCS Figure Dimensional Drawing of Arduino Mega 2560 Table : DCS Mass Summary Component Mass (lbs) Arduino Mega Adafruit 0 DOF IMU XBee Pro RPSMA MHz Duck Antenna SD Card Total.427 lbs 30

32 Observation Components Table Observation Component Alternative # Trade Study Alternative # Raspberry Pi 3 Model B with Raspberry Pi Camera Board v2 8 Mp, Add Battery Supply PROS: 080p at 30 fps Cheap ($60 plus a battery pack) Camera is small and largely adjustable Customizable Configuration Mass ~.76 oz w/o external battery and wiring Video has ability to be streamed Allows for use of other electronics CONS: Larger, 3.37 in x 2.22 in x.40 in (board only) Complicated setup This architecture has a small volume and allows for various camera configurations and customizability. It also is a low-cost solution. While, it also needs to be setup and programmed to do video and save to the SD card. Yet, there are plenty of resources for this setup. It also needs an external power setup. Finally, depending on quality of configuration it may be susceptible to failure in acceleration. 3

33 Table Observation Component Alternative #2 Trade Study Alternative #2 Arduino Uno and TTL Camera with SD breakout with Battery Shield PROS: Cheap ($70 plus a battery pack) Customizable Configuration Can be set to take photos will video taping Mass ~.06 oz w/o external battery and wiring 2.70 in x 2.0 in x 0.40 in (board only) Video has ability to be streamed Allows for use of other electronics CONS: Low Resolution (640x480 at 30 fps) Complicated setup This architecture has a small volume and allows for various camera configurations and customizability with added electronics and sensors. It needs to be manually setup and programmed to integrate it with the camera and SD card. It requires an external power source and wiring. Sensitive to failure during vibrations or large accelerations, if quality of setup isn t good. 32

34 Table Observation Component Alternative #3 Trade Study Alternative #3 GoPro Hero Session & Battery Supply PROS: Compact Simplified (Push button and go) Highly adjustable 080p at 60 fps On-board battery (Apprx. hour use).5 in x.43 in x.5 in (camera only) CONS: Expensive ($200 plus a battery pack) Not customizable Mass ~ 2.58 oz w/o external battery Rocket body protuberance or mirror system needed for video This architecture has the smallest volume, but is slightly heavier. It is the most expensive architecture. It is extremely easy to use and add an external battery. Yet, it is a unit that can t be customized or changed. Also, the whole unit is integrated as an all-in-one, therefore, it is harder to place or angle the camera. 33

35 Leading Alternative: Alternative # The Raspberry Pi camera observation subsystem deploys a Raspberry Pi 3 Model B and a Raspberry Pi Camera Board v2 to record video of the launch. It offers the most flexibility of the three systems with lots of resources available to setup the system in various configurations. It also allows for quick changes to be made over time allowing the system to develop into a more robust system over time. While, the volume is slightly larger than the other two alternatives it is still relatively small and there is plenty of room within the rocket body to house the hardware. This alternative is an inexpensive option and in addition, can provide quality video to verify the payload experiment and to get video of the entire flight. The camera being so small allows for it to be placed on the transition piece without creating a protuberance on the rocket body or utilizing a secondary mirror system to obtain the video. Figure Raspberry Pi Camera Design in Transition Piece 34

36 Figure Raspberry Pi Holder Dimensional Drawing Mass Summary Table Observation Subsystem Mass Summary Part Mass (oz) Raspberry Pi 3 Model B.59 Raspberry Pi Camera Board v mah Portable Commercial Battery 9.00 Additional Wires NA Total Mass

37 4.3.4 Launch Vehicle Leading Design and Mass Summary The leading design can be seen below in Figure in which the final dimensions, geometries, and leading designs for PDR were found. Figure shows the mass summary for the vehicle as well as the vehicle characteristics. Figure Launch Vehicle Leading Design 36

38 Figure Launch Vehicle Mass and Characteristics 37

39 4.4 Recovery Subsystem 4.4. Recovery Subsystem Overview The recovery system is being designed with a dual parachute deployment regime to satisfy requirement 2.. Once the rocket reaches apogee at approximately 5280 ft, the recovery bay altimeters trigger the aft parachute bay ejection charges and the cruciform drogue parachute will begin deployment. Once the drogue parachute has inflated, the rocket will ride the drogue until roughly 500 ft to ensure that the rocket does not drift more than 2250 ft. radially from the launch pad even in worst case condition winds. The forward bay altimeter will then trigger the main parachute bay ejection charge and the toroidal main parachute will begin deployment to ensure that requirement 2.3, a max kinetic energy of 75 ft-lbf on landing of each section is met. A general mission overview can be seen in Figure Figure Recovery System Overview 38

40 4.4.2 Main Parachute Recovery Component Trade Study Table Main Parachute Trade Study Alternative # Alternative #2 Alternative #3 FruityChutes elliptical parachute example Elliptical Parachute Elliptical shaped parachute with spill hole. PROS: Easy Construction Lower line tangle Small area CONS:.5 Cd FruityChutes toroidal parachute example Toroidal Parachute Toroidal shaped parachute, shape produced by extra set of suspension lines in center PROS: 2.2 Cd Smallest area CONS: Complicated Construction Higher tangle chance Heavenly hobbies hemispherical parachute example Hemispherical Parachute Hemispherical shaped parachute with spill hole. PROS: Easy Construction Low line tangle CONS:.5 Cd Highest area Easy construction allows for less chance of manufacturing error Low coefficient of drag compared to toroidal Takes up more volume than toroidal while packed Fewer suspension lines required than toroidal allow for less chance of line tangle Extremely efficient packing area due to low surface area required Very high Cd Complicated construction, possibly may avoid by obtaining commercial parachute Two sets of suspension lines increase line tangle chance on deployment Easy construction allows for less chance of line tangle Low coefficient of drag compared to toroidal parachutes Symmetrical shape produces more surface area, thus increasing packing volume Fewer suspension lines than toroidal allow for less chance of line tangle 39

41 Leading Alternative: Alternative #2 The current leading alternative design is the toroidal parachute for the main. This is due to its extremely efficient packing factor. The toroidal shape produces an extremely high Cd, and takes less surface area than the elliptical parachute. The only detriment of the toroidal, neglecting construction problems, is that the lines and parachute have a tendency to tangle. This is due to the nature of the parachute requiring two sets of suspension lines to create the toroid shape. To prevent this, an appropriate packing method will be used. The diameter of the parachute likely will be 72 with an inner diameter of 2.67 roughly, creating a projected area of ft 2. These numbers are per the commercially available Fruity Chutes Iris Super Compact 72 parachute Figure Dimension Drawing of Toroidal Parachute Mass Summary Table Main Parachute Mass Summary Item Mass (lb) Parachute and suspension lines 5.46 Rosco 500 lb swivel.06 Kevlar Shock Cord.76 40

42 4.4.3 Drogue Parachute Recovery Component Trade Study Table Drogue Parachute Trade Study Alternative # Alternative #2 Alternative #3 Shopify cruciform parachute example. Cruciform Parachute Parachute constructed of two rectangular gores. PROS: Extremely stable design Easy construction Light weight CONS:. Cd Subject to tangling FruityChutes elliptical parachute example Elliptical Parachute Elliptical Parachute, no spill hole. PROS: Easy construction.5 Cd CONS: Less Stable FruityChutes toroidal parachute example Toroidal Parachute Toroidal Parachute, shape achieved by an extra set of suspension lines on the spill hole. PROS: 2.2 Cd Low packing area CONS: Extremely difficult to achieve deployment at this scale Extremely stable parachute design is ideal for fragile material protection. Extremely simple construction. Tangling is common but does not affect stability. The elliptical parachute is marginally less stable than the cruciform parachute. Less likely to tangle than the other two alternatives. Overall a simple design with small deployment failure chance. This design should be avoided due to the difficulty of deployment at this scale. The interior suspension lines have a high tangle chance and deployment failure is extremely likely. The construction of the parachute at this scale will be very difficult. The parachute descent is less stable than the other alternatives. 4

43 Leading Alternative: Alternative # The leading alternative is the cruciform parachute. The main reason for the selection is due to its stability during descent. The fragile material protection experiment was the main factor taken in consideration making this choice. The parachute is subject to tangling, but the tangling does not greatly effect parachute performance due to the simple design. This parachute is also commonly used throughout egg drop competitions due to its stability. The parachute will have 40% of the main s area, 0.96 ft 2. To achieve this shape, the rectangular panels will both be 54 x8 producing an area of.25ft 2, the extra area adding safety margin insuring requirement 2.3 is met, with the largest separable section having a final kinetic energy of less than 75 ft-lbf. Figure Dimension Drawing of Cruciform Drogue Parachute Mass Summary Table Drogue Parachute Mass Summary Item Mass (lb) Parachute and suspension lines 2.73 Rosco 600 lb. swivel.06 Kevlar Shock Cord

44 4.4.4 Recovery Avionics Trade Study Table Recovery Altimeter Trade Study Alternative # Alternative #2 Alternative #3 Entacore AIM USB Device logs altitude and fires a charge to deploy recovery system PROS: Only requires a USB cable to interface with a computer Dual event programmable altimeter CONS: The unit cost twice as much as the stratologgercf The Entacore AIM USB system can perform the recovery needs, but costs nearly 00 dollars more than the stratologger system Missile Works RRC3 Device logs altitude and fires a charge to deploy recovery system PROS: Contains a third circuit that could perform other tasks (e.g. ejecting a payload at a specific altitude) CONS: Requires an interfacing system The Missile Works RRC3 can adequately perform recovery tasks and is fitted with a third firing circuit, however it is not foreseen that the third circuit will be utilized during the recovery segment. Perfectflite StratologgerCF Device logs altitude and fires a charge to deploy recovery system PROS: Simple/robust Cheapest unit CONS: Requires an interfacing system The Stratologger is the simplest altimeter system, and the cheapest. The product can perform the recovery tasks with minimal unused features and excess weight. 43

45 Leading Alternative: Alternative #3 The StratologgerCF is the leading alternative based on the trade study. The StratologgerCF in comparison to the alternatives is very basic. The StratologgerCF has only two firing circuits and can be purchased with extra pin-out circuit to attach an amplified speaker or LED. Of the three alternatives, the StratologgerCF is also the cheapest. In addition, it is the most cost effective alternative. These components will also be independent of the electrical circuits from the payload avionics. If there is a need of multiple altimeters, each one can also have a dedicated power supply satisfying Requirement RSR2.8. Mass Summary The StratologgerCF weighs oz. Figure Dimensional Drawing of Leading Design 44

46 Table Recovery GPS Trade Study Alternative # Alternative #2 Alternative #3 BRB900 GPS receiver and 900 MHz transmitter. PROS: System does not require license to operate CONS: Most expensive unit The BRB900 utilizes a 900 MHz transmitter with a range of 6-5 miles, significantly less than the TELEGPS, but does not require a special license to operate. Trakimo GPS receiver and gsm data transmitter PROS: System does not require license to operate Half the price of BRB900 and TELEGPS CONS: Needs cell service to transmit data The Trakimo system theoretically has unlimited range as long as its connected to a cellular tower. This form of data transmission isn t optimal since most launch site is southern California typically has bad cell reception. TELEGPS GPS receiver and 420 MHz transmitter PROS: Large operating distance CONS: Requires license to operate The TELEGPS has a long range that will not be fully utilized, also to use this system the operator has to have an amateur radio operators license. 45

47 Leading Alternative: Alternative # The BRB900 is the leading alternative for the GPS system for a couple reasons. The BRB900 is the most expensive system, but that is primarily due to the frequency and means of data transmission to a ground station. The TELEGPS uses a frequency of 420 MHz to transmit data. In the United States and many other countries this frequency requires a special license to use, like how a highpower rocketry certification is required to launch high power rockets. However, it is not too difficult to obtain this license, but the system should be able to be operated in the worst-case scenario. If the one person with the license cannot show up the GPS system cannot be utilized. The other alternative is the Trakimo. The Trakimo uses the cellular network to transmit its logged data to the internet, which in then accessible by virtually any device that has access to the internet. High power rocket launch facilities and fields tend be far from any city or cell tower, and therefore no means to transmit the data. In southern California, a fair amount of launch sites is nestled in-between mountains and hills which make for poor cell reception. This launch condition differs from Alabama s geography, but it is important that the team familiarizes themselves with how the system works during tests flights. The Trakimo won t allow the team the practice they need with the system. Although the BRB900 does not stand out in price and operational range it has the perfect balance of functionality and reliability that allows the team the most familiarity with the system in case it needs to be used. In addition, this ensures that the electronic tracking device will be fully functional during official flight day. Figure Dimensional Drawing of Leading Design Mass Summary The BRB900 weighs.763 oz with the battery. 46

48 4.4.5 Recovery Subsystem Redundancy To ensure that the recovery system is successful, a redundant deployment system will be employed in both the main and drogue parachute system. This will include an extra altimeter and black powder charge per parachute bay. These will be independent from each other to add redundancy to the system. The secondary system in each case will only be triggered if a faulty ejection is detected. This will be determined based off accelerometer data checked in both the main and recovery bays. If a faulty trajectory is detected, the secondary ejection will be triggered and parachute will deploy. If a nominal trajectory is detected, the system will remain idle Parachute Sizing and Safe Decent Analysis Per requirement 2.3, the maximum kinetic energy upon landing may be 75 ft-lbf. To ensure the requirement is met, the maximum velocities of individual rocket sections were calculated, shown below. Table : Kinetic Energy Analysis Component Mass Max Velocity (slugs) (ft/s) Nose Cone Forward Rocket Section Aft Rocket Section To find the needed area of the parachute, equation () was used below. A = 2W C D ρv 2 Eqn. This is just a rearranged version of the drag equation solving for area for the worst case using the following assumptions: ft/s, 27.9 lbs for rocket weight, 2.2 Cd for the main parachute, and slug/ft 3 for density. The area for solely the main parachute found from this equation is ft 2.The FruityChutes Ultra Compact Iris parachute has a total area of 48.8 ft 2 and an effective area of ft 2, fitting the requirement with some room to spare. The drogue parachute has an area of.25 ft 2 and a Cd of., also adding to the deceleration, and adding more margin for safety. As the design further develops, re analysis will be done with the mass to ensure the parachute selection still fits requirement

49 4.4.7 Recovery Subsystem Leading Components The recovery system consists of a dual parachute deployment system to satisfy requirement 2.. The main parachute is a toroidal parachute with a 72-inch outer diameter. Currently we plan to use the FruityChutes Ultra Compact Iris parachute. The parachute satisfies the area requirement to ensure that requirement 2.3 is met. The drogue parachute will be a cruciform parachute with an area of.25 ft 2 due to its stability. The parachute will also decelerate the rocket adding safety margin to the main parachute selection. Each of the parachutes will be attached to the rocket using ½ inch thick Kevlar cord rated to 500 lbs. The main parachute will be attached to the main parachute bay using a 500 lb Rosco swivel. The drogue parachute will be attached to the aft parachute bay using a 600 lb Rosco swivel Deployment Charge and Altimeter Layout At each end of the Recovery Bay there will be a small cup of black powder with an electric-match taped inside. The purpose of black powder charge is to create a large enough pressure inside the section the parachutes are housed, such that the sections will separate and release the parachutes. In Figure below, the ends Recovery Bay are annotated. Figure Annotated Recovery Bay To obtain an idea of the size of the charges, we calculate theoretical values. These values will later be validated through ground testing to ensure the recovery system is properly deployed. The estimated force needed to separate a six-inch diameter rocket is roughly 300 to 350 lbf. Since the charge can always be increased, the charge for the minimum value of 300 pounds has been calculated. The calculations begin with solving for the pressure the charge must create. Using, PA = F Where P = pressure, A = area, and F = force, We get, P = F/A Using F = 300lb and A = cross sectional area of rocket tube, we calculate: 48

50 300 lb P = π(3in) 2 P = 0.6 psi Next we use the ideal gas law to calculate the size of the charges: PV = nrt Where P = pressure, V = volume, n = mass, R = gas constant, T = absolute temperature. Using the following specific values, R = 266 in-lbf/lbm T = 3307 degrees R (combustion temperature) P = pressure in psi V = volume in cubic inches n = mass in grams. (Note: 454 g/lb) (Values taken from vernk.com) We come up with an equation for charge size as a function of length: n = (. 547 g in ) L The charges for the main chute and drogue chute are calculated separately because they do not have the same volume. The main chute compartment has a length of 8 inches, while the drogue compartment has a length of 4 inches: n main =.547 g 8in = 2.78 g in n drogue =.547 g 4in = 0.62 g in To ignite the charges electronic matches are placed with the charges and then connected to the respective ports on the altimeters. Each altimeter will have their own set of e-matches to ignite. Each altimeter will perform redundant ignitions in case one fails. The altimeters will be fastened to a sled that slides on to a set of all-threads that run through the electronics bay. Figure shows a simple representation of how the altimeters will be wired on the sled. The sled acts as a mounting point that is easily accessible and removable for ease of maintenance, primarily for replacing and charging the batteries. 49

51 Figure Recovery Bay Electrical Schematic Recovery Bay Overview Each module (separated sections) will be untethered to the device shall have a separate tracking device, however if the design goes as planned, all sections will be tethered to the main launch vehicle. The recovery avionics bay is separate from the payload bay. This will ensure that the electronics will not interfere with each other. The recovery avionics bay will house its own electronics such as altimeters, where it would be safe from all other onboard transmitting devices that can generate inadvertent excitation, magnetic waves, or anything else that can affect functionality. Each altimeter in the recovery system shall also be armed by an arming switch that can be accessed on the exterior of the rocket during launch configuration. Each of the arming switches shall be able to be locked and unlocked. Shear pins will be used for both the main parachute and the drogue parachute. The Recovery Bay will be constructed out of blue tube, with 7 inches in length and an outer and inner diameter of 6.08 and 5.97 inches, respectively, as shown in Figure The Recovery bay will be inserted into the lower section of the Main Parachute Bay and secured to the launch vehicle through a bulkhead with 4 #0-32 zinc-plated alloy steel flat-head screws placed 50

52 symmetrically around the circumference. The Recovery Bay will also serve as a coupler and allow section 2 to connect to section 3. The bulkheads on each end of the bay support a U-bolt, along with two charge canisters made from PVC pipe. The shock cords and shroud lines will connect to the launch vehicle via quick links and U-bolt. The U-bolt will be fastened to the bulkheads using epoxy for additional strength. The schematics of the Recovery Bay can be seeing in Figure The aft bulkhead is a major component in the launch vehicle and will experience most the impulse force from the main parachute deployment, thus will be designed to handle the loads from the deployment. To ensure the bulkhead and shock cord can withstand the impulse force several drop tests will be designed. Figure Front view of Recovery Bay 5

53 Figure Schematics of the Recovery Bay A new interior design for the recovery bay will be implemented to the launch vehicle. The main objective of the design is to ensure the safety of the avionics while allowing the insertion and accessibility of the avionics. In the previous design, there were problems when assembling the avionics sled inside the bay. The wires would get stuck in between the centering and sled, blocking the full entry inside the bay, this pose the possibility of damaging the hardware if a team member had to reach to correct the issue. The new design will have a permanent sled retainer made from 3D printed ABD material as shown in Figure which will be attached via screws to the bulkhead. Figure Sled retainer permanently attached to the bulkhead. The sled retainer will also serve as a guide for the avionics sled to be inserted into the recovery bay. It will allow the sled to stayed fixed and secure inside at all times. The installed avionics sled can be seen in Figure Also, two bulkheads will attach to the ends of the recovery bay and lock the Recovery bay, ready to be installed inside the launch vehicle. 52

54 4.5 Full Scale Launch Vehicle Interfaces and Integrations 53

55 4.5. Interface A: Avionics Integration Figure 4.5. Payload Control System and Observation System Payload control system and the observation systems are self-contained systems within the launch vehicle. The payload control system is designed to control the control surfaces based off a feedback loop with the IMU, Inertial Measurement Unit. The IMU feeds acceleration to the commercial RC airplane system to determine burnout and to ensure rotation is met and the initial rotation is matched after completing the two rolls. All data from the IMU will be processed on the commercial RC airplane system and saved on the external SD card. This is to help verify the rotation experiment requirements were met in flight. The Observation system uses the Raspberry Pi 3 Model B to do capture the video processing from the Raspberry Pi Camera board v2. While, the Raspberry Pi 3 Model B then saves the video to the external SD. This is a redundancy to verify that the rotation experiment requirements have been met in flight. 54

56 4.5.2 Interface B: Fin Integration Fin integration is a very important part of the rocket, and special care was taken to integrate the fins into the motor bay to insure maximum structural integrity. The design of the bulkheads must include accommodation for mounting. Figure shows the bulkhead design to integrate the fin in which the small hole in the center is the allocated space for the 54 mm L-class motor and casing. The large diameter will rest onto the inside of the body tube and be secured using counter-sunk screws from the outside of the body tube. This, along with the other bulkheads for the motor casing will allow the entire fin and motor bay to be taken out of the body tube to allow any necessary work to be done on the structure. The notches found on the fin bulkhead will allow the fin to be integrated into the motor bay using epoxy resin as a means of attaching the two pieces. Figure Fin Bulkhead Model Figure will be the basic design for the fin attaching to the bulkhead. It is important to note that the actual fin, airfoil, and geometry design probably will change. Therefore, the fin used in this model is an arbitrary fin design to show the integration of the fin to the motor bay. the fin body structure will extend down into a solid block with equidistant notches, and these notches will connect into the fin bulkhead with epoxy. The important aspects of this design are the interlocking design of the fin and bulkhead. This interlocking design will maintain a greater structural integrity opposed to fins epoxied to the outside of the body tube. Fins attached to the outside of the body tube have the ability to shear off the body tube during the max dynamic pressure that the fin will experience. An interlocking design will have the ability to both help in shear stress and offer added compression and tensile stresses during the rockets operation. 55

57 Figure Fin Integration Model Figure Shows the body tube configuration for this fin integration. This would be the greatest down side to using a fin integration because of the body tube has slits cut out to allow the motor bay to be taken out. As seen by the figure, the slit cut outs would damage the body tubes structure in the axial direction, allowing the chance for buckling, as well as any twisting moments that may act on the tube. With the motor bay installed into the body tube with a series of counter sunk screws it will be able to take much more stress and bulkheads for the fin and motor casing will add stiffness to the body tube, thus negating any ill-effects that slits might cause on the tube. Figure and Figure show all the pieces discussed above, integrated together in one section. As can be seen, the structure is very sturdy and integrates into the body tube flawlessly. Also, the manufacturing process to complete this design is simple and easy to carry out in a timely manner. 56

58 Figure Fin-body Tube slits Figure Side View of Complete Fin Integration 57

59 Figure Exploded View of Fin Integration Set-Up 58

60 4.5.3 Interface C: Launch Vehicle and Ground Station Figure Launch Vehicle Data Collection System and Ground Station The launch vehicle data collection system is designed to feed IMU from the launch vehicle to the ground station for a near live representation of the rocket on the GUI. This allows for a third redundancy in determining completion of the rotation requirement. While, CSV file will be created for later analysis after completion of flight. The ideal sampling rate for this system is 0 Hz from the launch vehicle XBee to the ground station XBee. 59

61 4.6 Mission Performance Predictions 4.6. Mission Performance Criteria Mission performance criteria, listed below in Table 4.6.-, describes how well the launch vehicle will perform beyond mission requirements based on the allowable range of values defined by the team. A high performance and effective launch vehicle will be characterized by a minimal difference between target and actual peak altitudes, optimal stability margin, minimal ground impact velocity, drift distance, completion of required rotations post motor burn-out, and adequate protection of the fragile materials. Table 4.6.-: Mission Performance Criteria Performance Criteria Description Goal/Allowable Range for Success Peak altitude Reach a target peak altitude of 5,280 feet AGL Minimize altitude difference from target peak altitude. Allowable range: ±75 feet Stability Margin Main Parachute Deployment Altitude Kinetic Energy upon ground impact The center of gravity must be located forward of the center of pressure to provide a stable flight. The main parachute must be deployed at an altitude of 500 ft. AGL. Each independent section of the launch vehicle must withstand maximum impact kinetic energy of 75 ft-lbf so that there will be no damage to the structure or any internal components. The CG and CP will be optimized so that static margin is in range: 2 caliber < SM < 3 calibers Have a redundant parachute deployment system which will ensure the parachute is deployed in range: ±50 feet Minimize the ground approach velocity: 0 ft/s < Velocity < 20 ft/s 60

62 Horizontal Drift Distance Launch Vehicle Ease of Assembly Recovery System Fire Safety RIS Ascent Stage Activation Launch Vehicle Structural Robustness The distance between the launch pad and each individual section must not exceed a drift distance of 2,500 feet. The launch vehicle shall be capable of being prepared for flight at the launch site within 4 hours, from the time the Federal Aviation Administration flight wavers opens. Both parachutes must be able to with stand fire hazard generated by activation of the ejection charges. The RIS must activate and complete two full rotations between motor burnout and apogee. The launch vehicle shall be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications. Minimize the distance: 0 ft. < Drift distance < 2,500 ft. Modular design of the launch vehicle sections, where assembly only requires fastening of each module with shear pins. Both parachutes will be wrapped in fire retardant Nomex blankets. RIS shall be able to sense motor burnout and autonomously activate the roll. The launch vehicle will be designed to be structurally robust with reasonable factors of safety. Recovery System Redundancy The recovery system must have redundant altimeters and ejection charges. The recovery system will be built with duplicate systems to prevent recovery system failure. 6

63 4.6.2 Flight Profile Simulations To verify that requirements are met by the current design, simulations using OpenRocket were used to analyze the flight profile. Trajectories were analyzed and 0, 5, 0, 5, and 20 mph. Figures to display the results of the varying simulations, in ascending order of velocity. Launch Rail Velocity (ft/s) Apogee (ft) Table MPH Flight Trajectory Results Optimum Delay (s) Max Velocity (ft/s) Max Acceleration (ft/s 2 ) Time to Apogee (s) Total Flight Time (s) Ground Hit Velocity(ft/s) Figure MPH Flight Trajectory Profile 62

64 Launch Rail Velocity (ft/s) Apogee (ft) Table MPH Flight Trajectory Results Optimum Delay (s) Max Velocity (ft/s) Max Acceleration (ft/s 2 ) Time to Apogee (s) Total Flight Time (s) Ground Hit Velocity(ft/s) Figure MPH Flight Trajectory Profile 63

65 Launch Rail Velocity (ft/s) Apogee (ft) Table MPH Flight Trajectory Results Optimum Delay (s) Max Velocity (ft/s) Max Acceleration (ft/s 2 ) Time to Apogee (s) Total Flight Time (s) Ground Hit Velocity(ft/s) Figure MPH Flight Trajectory Profile 64

66 Launch Rail Velocity (ft/s) Apogee (ft) Table MPH Flight Trajectory Results Optimum Delay (s) Max Velocity (ft/s) Max Acceleration (ft/s 2 ) Time to Apogee (s) Total Flight Time (s) Ground Hit Velocity(ft/s) Figure MPH Flight Trajectory Profile 65

67 Launch Rail Velocity (ft/s) Apogee (ft) Table MPH Flight Trajectory Results Optimum Delay (s) Max Velocity (ft/s) Max Acceleration (ft/s 2 ) Time to Apogee (s) Total Flight Time (s) Ground Hit Velocity(ft/s) Figure MPH Flight Trajectory Profile In summary, for each case, the landing velocity was below the velocity to meet requirement 2.3, landing with less than 75 ft-lbf on impact. The maximum altitude for each case was slightly above requirement., an altitude of 5280 ft. This was tolerable due to the simulation being an ideal flight, and altitude will be lower in actual flight than in simulation. The relative error will be determined in the scale model flights later this year. In addition, it was clearly seen that the wind velocity did not greatly impact the flight profile due to our high stability margin of 2.28 cal. One of the only concerns of simulation at this moment is a high vertical acceleration on flight, seen at launch. This will be investigated and compared to more hand calculations by the CDR. 66

68 4.6.3 Altitude Predictions The current state of altitude predictions is being done with OpenRocket software. The rocket parameters were set as close to current predictions as possible. The software currently predicts a maximum altitude of 5555 ft, allowing some safety for simulation assumptions for requirement 2. to be met. This will be compared to a MATLAB written script by the CDR Motor Thrust Curve The Aerotech L50 produces pounds of force and 3489 Newton-seconds of impulse. As illustrated in the curve below, the motor has a high spike at the very beginning of the burn. After the first grains burn out, the other grains allow the thrust curve to gradually decrease within approximately 4.6 seconds. This translates into a brisk launch rail clearance and a rapid acceleration toward the one-mile requirement followed by a thrust deceleration. With this performance, the motor can deliver a pound launch vehicle to one mile with sufficient margins. Peak Thrust lb Average Thrust 258. lb Burn Time 3.7 s Figure Aerotech L50P Thrust Curve 67

69 4.6.5 Launch Vehicle Load Verification For load verification, high stress points on the body will be analyzed. The first point to be looked at is the transition piece seen in Figure A buckling analysis was performed by hand on the piece since it will be hollow on the inside. An ultimate allowable buckling stress was found to be 3. ksi for the weakest material, which happened to be the ABS plastic used in 3-D printing. With the approximate weight above the transition piece being 40 lbf, an estimated acceleration of 284 ft/s 2 and a safety factor of 2, the approximate load the piece will need to withhold is 750 lbf. This load results in a stress of 200 psi, well below the ultimate buckling stress. The ultimate compressive strength of ABS was also analyzed. With an Fcu of 9.4 ksi the transition segment is well above the threshold for failure. Figure Model of Transition Piece, Hollow with 2 Thickness on Conical Section 68

70 Another high stress point on the rocket is the engine block, like design as Figure except with 6 brackets, specifically the shear tear out and bearing stress on the birch wood from the bolts. Assuming the force on the engine block is the thrust output of the motor, the force is about 270 lbf. Distributing this force across the six bolts in the engine block it comes out to 45 lbf per bolt. The bolts themselves will be long and approximately.026 in diameter. The engine block itself is.75, leaving.25 of length for the shear tear out through the birch wood. With these dimensions the shear stress is approximately 90 psi and the bearing stress is 730 psi. The FS for the shear stress is 2.0 and the bearing stress FS is 2.0, satisfying the minimum FS requirements for the rocket. The last stress point to look at is the recovery bay bulkhead, seen in Figure and The recovery bay will be experiencing a snatch load of 500 lbf. The shearing and bearing stress caused by the bolts and the washers on the U-bolt will need to be looked at. The snatch load will be distributed across eight bolts for the recovery bay, giving each bolt a load of 87.5 lbf. The dimensions of the bolts are long and.08 in diameter. This results in a shearing stress of 375 psi and a bearing stress of 740 psi. This results in a FS of 3.0 for the shear stress and a FS of 2.0 for the bearing stress. For the U-bolt, the force of 500 lbf will be distributed across two washers whose radius is.370. These dimensions result in a shearing stress of 430 psi and bearing stress of 740 psi. These results in an FS of 2.0 for the bearing stress and FS of 2.6 for the shearing stress. These FS all satisfy the requirement of 2.0 minimum FS. Figure Dimensional Drawing of Recovery Bay Bulkhead 69

71 4.6.6 Stability Margin From previous experience the OpenRocket values are credible and acceptable for estimations for PDR calculations. The stability margin is measured in calibers, defined as the distance between the CG and CP, as depicted by the red circle in Figure One Caliber is the maximum body diameter of the launch vehicle. As such, the stability of the launch vehicle is the ability to stay in flight and point in the upwards direction without tumbling or spinning out of control. An overstable launch vehicle means that it has a greater tendency to weathercock if there is any wind during launch. As the launch vehicle flights, upward it will gradually turn into the wind, making the launch vehicle drift more than its intended. In typical model rocketry, a stability margin of to 2 calibers is appropriate if the length-to-diameter ratios are approximately 0. The length-todiameter ratio for Architecture and 2 are 5.4 and 4.6 respectively. Thus, a larger stability margin is desirable to keep the launch vehicle stable. A range of values between 2.0 and 2.5 calibers are considered acceptable by the team. As shown in Figure the stability of margin for Architecture is 2.38 calibers, and the stability for Architecture 2 as shown in Figure B is 2.28 calibers. Both values are within range, thus the launch vehicles are deemed stable. Once the final Architecture is selected, all changes to the weight of the launch vehicle will be closely monitored and updated in the OpenRocket model to ensure that the stability margin does not exceed 2.5 calibers nor fall below one diameter of the launch vehicle. Figure OpenRocket center of gravity and center of pressure locations for Architecture Figure OpenRocket center of gravity and center of pressure locations for Architecture-2 70

72 4.6.7 Calculated CP and CG In order predict how the launch vehicle will behave during flight, a careful consideration when designing the launch vehicle was emphasized to provide adequate stability. We say the launch vehicle is stable when it is designed to counteract any rotating forces caused by the wind or other disturbance that can cause the rocket to tilt in an undesired orientation. With these criteria, we model the launch vehicle to fly at a small angle of attack, to ensure the stability and improve the launch vehicle's performance. Center of Gravity: For the purpose of the PDR, the center of gravity (CG) was calculated using the OpenRocket, which calculates locations of the CG and CP. The team will narrow down the two Architectures being considered, based on performance with the help of the OpenRocket. When modeling the launch vehicle, the distribution of masses and proper location of the main components as well as the total length was analyzed to ensure the center of gravity is ahead of the center of pressure. If the CG needs to shift up for stability purposes, the nose will be left hollow to allow extra space for weight. The CG is crucial to calculate because it is the balance point of the launch vehicle. To calculate the CG, we chose the tip of the nose cone as the reference line, taken the centroid about this reference line and multiply by its respective weight gives a moment. The moments are summed and divided by the total weight of the launch vehicle. Figure gives a visual of where the reference line is and the distance of Architecture- CG which is inches from the reference line. Figure shows Architure-2 to have a CG of 53.8 inches. Center of pressure: The center of pressure involves air pressure forces acting on the rocket during flight. The CP of the rocket is the point at which all the air pressure forces concentrated at a single point. When designing the launch vehicle careful consideration of the fin and nose cone design was considered to keep the CP closer to the tail end. Placing the fins at the end of the rocket moves the CP closer towards the tail end, which helps increase stability while increasing the nose cone raises the CP closer to the CG. For the purpose of the PDR, we utilized OpenRocket to calculate the center of pressure for both Architectures. As shown in Figure the CP for Architecture- is 69.28, again using the nose cone as the reference line. For Architecture-2 the CP is Both designs show that the CP is indeed aft of the CG, confirming the stability of the launch vehicle to be within the range of desired values. Table CG and CP Values for Architecture and 2 Center of Gravity Center of pressure Architecture in in Architecture in 67.5 in 7

73 4.6.8 Kinetic Energy To ensure that requirement 2.3 is met, less than 75 ft-lbf during landing, the recovery system will be designed to ensure proper impact velocity occurs during landing of each individual section using equation. ke = 2 mv2 Eqn. Table x.x Kinetic Energy Analysis Component Mass Max Velocity (slugs) (ft/s) Nose Cone Forward Rocket Section Aft Rocket Section To meet requirement 2.3, the sections will need to land at a maximum velocity of ft/s Designing to a landing velocity of roughly 3. ft/s with currently accessible parachutes produces a factor of safety of roughly

74 4.6.9 Drift Calculations (no wind, 5 mph, 0, 5, 20) For the calculations OpenRocket was used to calculated the lateral distance of the rocket at varying wind speeds of 0, 5, 0, 5, and 20 mph, directly perpendicular to the launch vehicle. Figure shows a visual representation of the simulated range mph 5 mph 0 mph 5 mph 20 mph Max Drift Distance Figure Simulated drift distances at varying wind velocities With the current predicted parachute dimensions, the drift is simulated to pass in every wind case. During the CDR phase, a MATLAB script will be written to verify results and confirm the current deployment altitude of 500 ft will suffice to meet the 2250 ft max drift radius requirement. 73

75 5.0 Preliminary Launch Plan Overview 5. Test Launch, Pre-Launch and Launch Compliance Checklist Vehicle shall carry one commercial available barometric altimeter for official altitude determination o This altimeter shall report altitude via series of beeps At LRR, the NASA marked altimeter will be used for official scoring The team shall report to the NASA official the altitude with the officially marked altimeter The rocket must be flown at competition launch site The rocket must have proved to be reusable and recoverable before flight day Launch vehicle shall be prepared within 4 hours before launch Launch vehicle must be in ready configuration at a minimum of hour Launch vehicle shall have no external circuitry Launch vehicle shall be launched with a 2 Volt direct current firing system After completing full scale demonstration flights, no components shall not be modified Vehicle shall be in fully ballasted configuration during full scale test Full scale motor does not have to be flown during full scale test Successfully recover subscale model Vehicle and recovery system shall be functioned as designed Payload does not have to be flown but vehicle must have a simulated mass All teams shall use launch pad provided Must implement EIT Accessibility Standards 5.2 Preliminary Checklist of Final Assembly and Launch Procedures 5.2. Recovery Preparation Powder Charge Preparation:. Ensure handler and those in the vicinity are wearing safety glasses. 2. Insert the e-match into the modified shotgun shell. 3. Ensure seal at insertion point. 4. Carefully pour the measured amount of 4F black powder into the modified shotgun shell. 5. Pack the remaining space of the modified shotgun shell with dog barf wadding. 6. Seal the top of the modified shotgun shell with blue painter s tape. 7. Place prepared powder charge into the ammunition can until ready to mount. Recovery Bay Preparation:. Perform visual inspection of all electronics and wire connections. 2. Ensure handler and those in the vicinity are wearing safety glasses. 74

76 3. Ensuring powder charges are facing away from all personnel; connect main powder charges to exterior terminals on the fore payload bay bulkhead. 4. Mount powder charges onto the payload bulkhead. 5. Ensure continuity and secure placement of powder charges. 6. Bolt bulkhead and main parachute bay onto the front of the payload bay. 7. Ensuring powder charges are facing away from all personnel; connect drogue powder charges to exterior terminals on the forward recovery bay bulkhead. 8. Mount powder charges onto the front bulkhead. 9. Ensure continuity and secure placement of powder charges. 0. Connect altimeters to the terminal leads; Ensure continuity.. Announce the intention to connect batteries and clear area of all unnecessary personnel. 2. Connect two (2) batteries. 3. Carefully slide electronics board into place. 4. Bolt bulkhead and drogue parachute bay onto the front of the recovery bay. Parachute Preparation:. Perform visual inspection of nylon shock cords. 2. Perform visual inspection of Nomex Thermal Protection Blankets. 3. Perform visual inspection of connection points (quick links and eye bolts). 4. Perform visual inspection of the parachute. 5. Attach Nomex Thermal Protection Blankets to the parachute/shock cord connection point. 6. Fold parachute according to the proper folding procedure. 7. Wrap the folded parachute in Nomex Thermal Protection Blankets ensuring there is no exposed parachute material. 8. Connect to the respective eye bolt on the recovery bay. 9. Insert into the respective parachute bay. 0. Set aside until ready to mount Scientific Payload Preparation RIS Preparation:. Perform visual inspection of all electronics, wire connections, and mechanisms of the Rotation Induction System. 2. Connect batteries to the RIS 3. Activate RIS 4. Run system check and perform visual confirmation of active mechanisms. FMP Preparation:. Perform visual inspection of 3D printed pill and all mechanisms of the Fragile Material Protection payload. 2. Open the pill and insert the fragile material. 3. Seal the top of the pill. 4. Insert the pill into the tubing suspension. 5. Close the FMP payload bay. 75

77 5.2.3 Motor Preparation Motor Assembly:. Ensure handler and those in the vicinity are wearing safety glasses. 2. Ensure motor casing not damaged or modified. 3. Unwrap the motor and place on an appropriate surface. 4. Ensure all materials listed in the manual are present and not damaged. 5. Apply a thin film of silicon O-ring lubricant to the inside of the motor casing. 6. Apply a thin film of silicon O-ring lubricant to the outside of the motor. 7. With the protective nozzle cap on, insert the motor into the motor tube. 8. Apply lubricant to the threads of the aft closure. 9. Remove the nozzle cap and thread aft closure onto the case. Tighten until the motor is properly seated. 0. Reinstall the nozzle cap onto the nozzle.. Wipe clean the motor casing ensuring there is no residue. 2. Insert the motor casing into the motor mount. 3. Attach retention ring. 4. Insert motor mount into the motor bay. 5. Bolt into place Final Assembly and Launch Preparations Final Assembly:. Connect the Nose Cone to the main parachute bay with shear pins. 2. Connect the front of the Payload Bay to the drogue parachute bay with shear pins. Setup on Launcher:. Lower launcher to the horizontal position. 2. Ensuring no personnel are in the flight path of the launch vehicle, carefully slide the launch vehicle onto the launch rail. 3. Ensure the launch vehicle is properly seated on the launch rail. 4. Set Payload Bay door to the open position. 5. Ensure the igniter is properly fed into the Ignition Insertion System. Launch Procedure: (unnecessary personnel removed from the area). Once the launch vehicle is in launch position and the igniter is inserted, arm the electronics. 2. Safety officer check to ensure the checklist is properly completed. 3. The LCO enables the master arming switch. 4. Once LCO allows, the hard switch will be activated. 5. The LCO will commence the countdown of 5 seconds. 6. Once the countdown is completed. The LCO says fire and ignition is triggered. Post-Flight Inspection:. Visually track launch vehicle and payload from the time of launch to the time of recovery. 76

78 2. Assemble a team of two groups of at least two team members to recover the launch vehicle and the payload capsule. 3. Wait a minimum of sixty seconds before securing the launch vehicle and payload capsule. 4. Inspect the launch vehicle s external components for any clear signs of damage. 5. Document the launch vehicle through inspection and photographs for the later assessment. 6. Download video data and review altimeter data. 77

79 6.0 Payload Criteria 6. Primary Payload Experiment 6.. Payload Objective and Experiment Our primary payload objective shall be the successful performance of launch vehicle roll and counter roll as described by NSL 207 Handbook Experiment Requirements, Option #2 (req. 3.3): Teams shall design a system capable of controlling launch vehicle roll post motor burnout. o The systems shall first induce at least two rotations around the roll axis of the launch vehicle. o After the system has induced two rotations, it must induce a counter rolling moment to halt all rolling motion for the remainder of launch vehicle ascent. o Teams shall provide proof of controlled roll and successful counter roll Teams shall not intentionally design a launch vehicle with a fixed geometry that can create a passive roll effect Teams shall only use mechanical devices for rolling procedures Success Criteria The Roll Induction System (RIS) shall use the following as criteria for a successful experiment: Requirement Number Success Criteria 3.3. RIS controls roll action of launch vehicle during flight Following motor burnout, the RIS rolls the rocket at least two complete revolutions. The rocket s roll axis position at the time of motor burnout shall be used as the point of reference Following the roll maneuvers, the RIS initiates a counter roll maneuver to halt all further roll motion for the remainder of ascent Data telemetry received by the ground station shall be used to confirm successful roll and counter roll maneuvers. Recorded video from the rocket s Observation Bay shall also serve as confirmation after post-recovery retrieval All rolling motion shall be the result of active operation from the RIS Only mechanical devices shall be utilized in operation of the RIS. Systems utilizing items such as pressure vessels, chemical energy devices, etc. will be avoided. 78

80 6..3 System Trade Study Table Primary Payload System Trade Study - Alternative RIS A (Inertia Flywheel Design) Brief Description: This is a design for an inertia flywheel in order to accomplish the task for the Roll Induction System. It harnesses the physics of moments of inertia in which a flywheel, large thick walled cylindrical metal tube, which has a large moment of inertia is given an angular acceleration using an electric motor. This electric motor will induce a torque and due to equal and opposite moments on the rocket, the equal torque will act on the rocket as a whole and induce an angular acceleration thus spinning the rocket in flight. Pros: There are no aerodynamic protrusions outside the body tube, thus leading to a more aerodynamically stable rocket This has the ability to be quick and responsive when given a command Since the moment of inertia of the flywheel is internally in the rocket, it is not effected by speed, wind, or other conditions found on the outside of the rocket A critical failure of this design does not lead to a critical failure of the rocket Cons: Very Heavy design; creating a strain on the rocket s overall weight Very expensive design; creating a strain on the budget Demands a lot of power; more money for batteries and bigger heavier batteries Complicated Design full of potential issues High acceleration forces pose a problem for the motors A heavy large metal cylinder on a rocket during flight could cause safety concerns The Inertia Flywheel Design is a potential candidate for the roll induction system, however it could be more complicated and more trouble than it s worth. The complications that come from this design are high in number and another design could be more efficient. NASA uses a similar design to this with satellites called a Reaction wheel, and the reason why this is used in space is because there are no aerodynamic forces in a vacuum and this method is one of the only few ways of rotating an object in space. Given our rocket is still in the atmosphere, and can utilize the aerodynamic forces of air while climbing, it would be more efficient to use another design choice, this design might work extremely well, but given the limited resources we have there might be a better solution. 79

81 Table Primary Payload System Trade Study - Alternative 2 RIS B (Fin Aileron Design) Brief Description: This approach to the roll induction experiment utilizes a more conventional flight surface concept. Servo actuated ailerons located at the trailing edges of the rocket s fins will employ aerodynamic forces to roll, counter roll, and stabilize the launch vehicle along its roll axis during flight. Each of the rocket s four fins will have an identical aileron configuration, while the root chord of the fins will remain stationary and fixed to the body tube. Pros: Low response time; quick actuation of roll and counter roll Minimal mass required, especially compared to inertial flywheel concept Minimal power consumption from launch vehicle s batteries Conventional approach to roll control; more data and general information available for use Cons: Adds a degree of fragility to fins; increases chances of fin structural failure upon recovery touch down Needs refined and durable mechanical design Difficult to simulate realistic conditions for testing Improper operation could lead to flight trajectory instability The fin aileron control surface concept utilizes the low altitude atmospheric flight profile to its advantage. It s an energy efficient approach since it does not require substantial additional mass to achieve the payload experiment success criteria. The lighter overall weight of the rocket and less powerful engine required also make it a safer choice. It is not without its risks though. Improper operation of the control surfaces could easily lead to flight instability, and possibly even loss of the launch vehicle. It is also a difficult and risky design to test. These disadvantages will need to be weighed accordingly during consideration of all roll induction options. 80

82 Table Primary Payload System Trade Study - Alternative 3 RIS C (Open-Closed Fin Stabilizer Design) Brief Description: The objective of this design is to deploy an active system that can quickly stabilize and induce a counter roll despite cross-winds and other disturbances on the launch vehicle. By having a second set of fins the system has full control of the roll axes while aiding in the stabilization of the launch vehicle. After post motor burnout, two linear actuators will be utilized to deploy the fins out of the airframe. The control system will then require attitude determination about two stable axes, to completely stop roll. Afterwards the system will determine if it s necessary to adjust the fin angle to induce a counter roll or second, have an electric motor give an angular acceleration on the fins to induce a torque and allow the launch vehicle to rotated. Pros: There are no aerodynamic protrusions outside the airframe during flight burnout, which will lead to a more aerodynamically stable rocket The ability to be quick and responsive when given a command The ability for fins close back in before landing will help secure reusability The system can aid in stabilizing the rocket about the pitch and yaw axes Linear actuators will reduce the amount of external power necessary to push out fins and help keep cost down Cons: The system can potentially act as an airbrake system Actuators not strong enough to resist aerodynamic loads, and can possibly buckle One or two fins breaking off during deployment due to aerodynamic forces, can cause the launch vehicle to turn into the wind and increase drift distance Size of linear actuator can post a dimension constraint, having to increase the size of the airframe to 8 inch. High acceleration forces pose a problem for the motor The Open-Close Fin Stabilizer design is a potential candidate for the roll induction system. It must be capable to achieve roll induction, while making the design and construction feasible for the team. The main advantage of having a separate fin system is that you keep the rear fins fix, which insures you don t disrupt the aerodynamics of the launch vehicle during motor burnout. Also, the launch vehicle will not rely on a single system for stability and achievement roll induction. The main issues will be making sure the fins deploy properly and withstand the aerodynamic loads, if a fin breaks the rocket can drift in an undesirable direction. Proper testing will be done mitigate these issues. 8

83 6..4 System Leading Design For this payload challenge, the system leading design candidate is best chosen through process of elimination. RIS A presents serious design and safety issues. Given the large mass of the momentum wheel, structural reinforcement of the launch vehicle and the actual mechanism itself adds several layers of complexity. A bearing system would have to be developed that would allow the free rotation of the flywheel while the rocket experiences the large acceleration of engine burn. The momentum wheel would have to be at least 0lb to generate sufficient torque to rotate the rocket in the time allotted. In addition to the large mass of the system itself, large batteries would also be required to keep the system running on the launch pad for an hour pursuant to Req..8. The large overall mass increase to the rocket would require a much larger motor, and presents safety issues in the event of recovery failure. RIS C is an interesting hybrid concept; however, we don t feel its technology readiness will allow us to quickly fabricate a prototype and begin testing. We have thus concluded that RIS C, the Fin Aileron Control Surface, is the best option for the Roll Induction Experiment. It utilizes the low altitude, atmospheric flight profile of our mission while minimizing the mass burden on the launch vehicle. The design of the underlying mechanical system will pose a challenge, but it is one we feel capable of meeting and exceeding. We envision a coupled mechanical system that will constrain the control surfaces to move together, to prevent any other trajectory change besides our intended roll. Thus, a coupled system will mitigate safety concerns regarding errant trajectories Leading Mass Summary Table Mass Summary Component Material Number Mass (Wt. lb) Fin Plywood w/fiberglass 4.78 Coating Aileron Plywood w/fiberglass 4.08 Coating Aileron Axle Stainless Steel 4.28 Servo Plastic, metal Leading Drawings and Electrical Schematics Total.33 82

84 Figure Preliminary Control Surface Fin Assembly, Sections -3 Figure : Preliminary Fin Design, Section 83

85 Figure : Preliminary Fin Design, Section 2 Figure : Preliminary Fin Design, Section 3 84

86 Figure Electrical Schematic 85

87 6..7 Interfaces: Payload and Launch Vehicle The payload leading design will be the RIS B controllable fin surface design which was discussed in detail in section This system will be heavily involved in the fin integration design of the rocket in which careful consideration will be taken for a variety of design criteria. These design criteria will be: temperature control within the body tube, stress concentration of the fin integration connections, volume allocated for mechanical servos, and fin airfoil design. Temperature Control can play an important role in the design of the motor bay in which the payload mechanics and electronics will be located. Thermal calculations have not yet been done for the motor casing but we can assume with some level of certainty that the thermal temperature could be a problem with the electronics for the RIS. This temperature could be alleviated using a thermal insulation on the motor casing such as Kevlar that would keep the electronics insulated from the high temperatures of the motor casing. This could or not be a problem but it would be important to note this and to add alternative designs to alleviate the temperature if necessary. The current fin integration will lock into the motor bay bulkheads and allow the fins to be stable during the high stresses of aerodynamic drag, as well as during high rocket acceleration at lift off. This interlocking fin-bulkhead design could be problematic for the electronics of the RIS. The controllable surface on the fin will run an axil inside the fin down through the body tube and ending near the motor casing. Giving this design the fin integration must be strong enough to survive flight while also allowing wires, servos, and circuits to be mounted to the bulkheads. Cutting wire holes or mounting hardware for these electronic components can lead to a structural compromise on the fin integration and thus a failure of the rocket. As of right now we know the servos for controlling the actuating surface on the fin must be located within this section and be mounted to this section. Calculations must be completed to start iteration process to both improve and optimize the design for the rocket. The volume within this section can be a tight fit and could pose a problem for the necessary electronics. Enough room must be allocated for the motor casing which is 2.2 in diameter and outer body tube diameter of 6 in. This will leave a small amount of space allocated for internal components which could lead to problems with manufacturing and integration. Currently we are trying to implement an airfoil design into the fins, the purpose of this is too decrease drag as much as possible, make the manufacturing process easier, and increase the stability by having symmetric and equivalent sized fins. By picking an airfoil we can decrease the drag by allowing a smoother path for the flow to travel around the fins. Also, 3D printing the fins with fiberglass reinforcement can make the manufacturing process more systematic and easy to carry out. The fins will be manufactured by 3D material and due to this the accuracy of each fin 86

88 will be within microns which decreases the chances of manufacturing errors leading to differences of aerodynamic forces. The entire point of the payload is too roll the rocket; it would be preferable to keep the rocket as stable as possible on ascent so the RIS can function without unnecessary aerodynamic impediments Precision, Repeatability, Recovery Precision The precision of the system relies heavily on three key factors the fidelity of the accelerometer, accuracy of servos, and the iteration rate of computing hardware. The precision of the system will rely on the weakest link of the three. Based on current research of the hardware, the controlling hardware maybe the root of the bottleneck. This specific part of the system must receive sensor data, interpret it, and then send an output to the servo. The system will need to be optimized for the best refresh rate and accuracy. Repeatability Barring any extreme unforeseen scenario, the systems performance should be completely repeatable. The system, as designed, does not have any single use hardware. Furthermore, the system is designed to be as robust as possible since components will not be readily accessible once the system is manufactured. However, to continue proper function and repeatability the battery should be recharged after every flight. Recovery The system is integrated into the rockets airframe. Upon descent and touch down the system will be recovered with the rocket. 87

89 6.2 Secondary Payload Experiment 6.2. Payload Objective and Experiment The Fragile Material Protection, FMP, is the secondary payload in our rocket. The objective of this payload is to protect a fragile object, one or more, from the forces of the launch itself, and the impact of the rocket landing. This specific item(s) will also be unknown to the team members until launch day. The specifics of the FMP payload from the NSL University Handbook are as follows: The object(s) in need of protection will be able to fit in a cylinder container 3.5 in diameter and 6 in height. The object(s) will have a max total weight of 4oz. With that knowledge, the payload is being designed with the intention of decreasing any acceleration experienced by the object(s) to prevent its breaking Success Criteria Our standards of success for this experiment are the following which access the requirements for the secondary payload: Ensure fragile materials survive acceleration duration takeoff Ensure fragile materials survive parachute deployment Ensure fragile materials survive landing Reusability Replaceable 88

90 6.2.3 System Trade Study Table Secondary Payload System Trade Study - Alternative Alternative # Surgical Tubing Suspension Plastic "Pill" surrounded in surgical tubing under tension that will resist acceleration. Wooden bands with the surgical tubing running through will secure the pill in both the x,y,z axis. PROS: Flexible Low Cost Lightweight Easy to fix and modify CONS: Spacious Complexity While somewhat complex to manufacture, this design will have many benefits on the day of the launch. The entirety of this system will be able to be removed from the rocket with ease as it will be clasped to and internal wall of the rocket. This will allow for quick insertion and removal of the material. Also with numerous surgical tubes, a failure of one or two does not hinder its performance fatally. The tensions of the tubing can all be adjusted at any time and this will allow for quick adjustments if needed. 89

91 Table Secondary Payload System Trade Study - Alternative 2 Alternative #2 Air Bag Bags filled with air, enveloping fragile material placed above and below the object. Filling the bags with air to specific pressures will result in different resistance to acceleration applied to the fragile material. PROS: Most Simple Lightweight Low cost CONS: Spacious Little to no back up Little durability Although simple and low cost to build, air bags have no room for a back-up protection system. In the case of a deflated bag, there is no other cushioning nor safety system to be able to keep the material safe. Securing the payload in this system to the rest of the rocket is difficult as there are no attachment points anywhere on the airbags. Also, we would need additional tools, a pump to inflate the bags, to be brought out to the rocket when placing the material inside. 90

92 Table Secondary Payload System Trade Study - Alternative 3 Alternative #3 Box Suspension Padded box with fragile material inside suspended in the rocket body using springs attached to the corners of the box and to the walls of the rocket body. PROS: Structural integrity CONS: Spacious Heavy Greatest Complexity Difficult to install Even though this design provides plenty of shock absorption and structural integrity it is complex enough to allow for mistakes to occur when setting up on the launch day. This design also requires more than 8" of the rocket in height, the most required out of all of our designs. This will also be the heaviest design out of all three choices. 9

93 6.2.4 System Leading Design Leading Alternative: Alternative # Analyzing the three architectures for the FMP, we deduced the Surgical Tubing would be the best alternative. Compared to the Box Suspension, the Surgical Tubes allow for an overall optimization of space and weight inside the rocket. Also, unlike the Airbag, this architecture would be most likely to survive the flight of the rocket making it ready for re-departure with little adjustment after each flight, making it extremely reusable. A 3D rendering of Alternative # can be seen below in Figure Figure Leading Alternative sans suspension system FMP Mass Summary Table Mass Summary Item Purpose Mass(lbs) Surgical Tubing Holding the FMP pill in place 0.2 White Printer Filament 3-D Printed Pill to hold fragile materials ~0.3 Egg Crate Foam Reduce stress on fragile materials ~0.03 Plywood Maintain shape of payload and hold tubing ~0.73 Sponge Act as a cushion in case tubing extents too far ~0.2 Total ~.47 92

94 6.2.6 Drawings and Schematics Figure and Figure are dimensional drawings and descriptions of our chosen alternative, alternative # Surgical Tubing Suspension. Figure Dimension Drawing of Leading Design Figure Descriptive Drawing of Leading Design's Suspension System 93

95 6.2.7 Interfaces: Payload and Launch Vehicle The Fragile Material Protection payload will be located in the green cell in Figure below. It needs no connections to any other part of the rocket so it can be totally self-contained. Above the Fragile Material Protection will be a separation point of the rocket, this will allow the removal of the FMP from the rocket. The other side will be a solid wall where the FMP will be attached with a series of clasps that will allow easy removal. Easy removal is necessary because the material in need of protection will be given just before launch and we cannot waste time dismantling the rocket to place the material inside. So being able to quickly remove the entire FMP from the rocket and placing the material in the correct manner is a necessity for the mission s success. Figure Side View of Rocket Precision, Repeatability, Recovery The Fragile Material Protection does not take any precise measurements, and does not include any instruments measuring quantitative values, other than the object breaking or not, as its entire purpose is to be a protective layer for the fragile payload. However, the FMP payload design is completely reusable assuming no damage are obtained during the launch. In the case of damages to the FMP, all parts, including the inner capsule can be replaced easily. The FMP can function well even with a few broken surgical tubes, and can be replaced quickly to perform optimally again. Our FMP design does not have its own recovery system, rather, it relies on the rocket's drogue and main parachutes. In addition, the FMP can be retrieved from the rocket without the destruction or removal of any other part of the rocket and can be reinstalled again separately with no alterations, making repeatability of our design extremely easy. 94

96 6.3 Payload Standard Preliminary Operating Procedure Figure 6.4- Payload Standard Preliminary Operating Procedure Flowchart 95

97 7.0 Safety 7. Safety Officer Identified Responsibilities Safety Officer, Aerodynamics Engineer Michael Nguyen Safety Officer Responsibilities. Generate a launch and safety checklist by FRR submission and used during launch days and LRR. 2. Filter out unsafe designs of the launch vehicle and payload experiments 3. Compliance of team members to wear safety equipment and follow procedures carefully during manufacturing 4. Monitor team activities during: a. design of vehicle and launcher b. construction of vehicle and launcher c. assembly of vehicle and launcher d. ground testing of vehicle and launcher e. recovery testing f. sub-scale testing g. full scale testing h. launch day i. educational engagement activities 5. Implement procedures for construction, assembly, launch, and recovery activities 6. Manage, maintain and write the team s hazard analyses, failure mode analyses, procedures for the overall launch vehicle system and the payload experiment 7. Attain and record all MSDS of materials used during construction and assembly 8. Write and enforce document stating that the team will abide by the rules of the FAA 9. Write and enforce document stating that the team will abide by local rocketry club s RSO during test flights 0. Determine all the risks and delay impacts and analyze the means of mitigation. Determine environmental concerns and effects of the launch vehicle system and payload 2. Create a Safety Compliance Document to fulfill Requirement SR4.5 and SR4.6 96

98 7.2 Risk and Delay Impact There are other potential risks, aside from the actual launch vehicle itself, to the project. These risks, including delays in report and budget deadlines, can be detrimental and are covered in Table Deadlines/Budget Risk Assessment Table 7.2-: Deadlines and Budget Risk Assessment Hazard Cause Effect Pre-Mitigation RAC Failure to meet Proposal deadline Failure to meet PDR deadline Failure to meet CDR deadline Incomplete safety plan Incomplete educational outreach plan Incomplete technical details Incomplete payload design Inadequate subsystem design Launch vehicle design that does not meet functional requirements Unacceptable payload integration Unsuccessful launch of sub-scale launch vehicle Insufficient maturity in design since PDR Proposal not accepted by officials Unable to pass PDR review with go ahead to manufacture Unable to pass CDR Review with go ahead to test launch full-scale launch vehicle Pre-Risk Mitigation Post-Mitigation RAC 3E E Low Weekly team meetings Systematic approach to distribution of work Immediate response to any activity that does not move the team forward E Low Well thought out approach to review preparation Complete organized launch vehicle and subsystem design Frequent review of requirements to ensure positive design progress D Moderate Implement systems engineering techniques to organize launch vehicle and payload design and keep project on schedule 3E 3E Post- Risk Low Low Low 97

99 Failure to meet FRR deadline Failure to receive necessary project funding Unacceptable final launch vehicle design Unable to demonstrate payload completeness and correctness via video Failure to demonstrate successful full-scale launch vehicle launch Failure to present acceptable testing of recovery system and interface with ground system Not enough fundraising Not enough community outreach and support requests Unable to pass CDR with go ahead to compete in final launch Unable to purchase necessary materials and equipment Insufficient traveling funds Constant review of requirements to ensure they are being met by design components Analysis and testing of key features of payload experiments Weekly meetings D Moderate Complete analysis on critical aerodynamic parameters during flight Top to bottom testing of necessary code for ground station and avionics Complete and thorough analysis and testing of recovery system including parachute sizing and material selection C High Create a well-designed and thought-out funding plan Develop a welcome package that can be distributed to local companies requesting support Reach out to college grants and programs for support 3E 2E Low Low 98

100 7.3 Failure Modes and Effects Analysis (FMEA) The risk matrix will be utilized to categorize and mitigate potential hazards involved with the project. Two factors, likelihood and severity, accompanies each risk. Tables 7.3- and elaborates on how the likelihood and severity factors are given. These values are used to create the Risk Assessment Codes (RAC) which determines the risk of each hazard. Tables and illustrates a color-coded chart for the RAC and corresponding risk level. Table 7.3-: Likelihood Definitions Description Qualitative Definition Quantitative Definition A - Frequent High likelihood to occur immediately or continuously Probability > 0.9 B - Probable Likely to frequently occur 0.9 Probability > 0.5 C - Occasional Expected to occur occasionally 0.5 Probability > 0. D - Remote Unlikely to occur but reasonable to expect occurrence at some point in time 0. Probability >0.0 E - Improbable Very unlikely to occur with no expect occurrence over time 0.0 Probability Description - Catastrophic Personnel Safety and Health Loss of Life or permanent injury Table 7.3-2: Severity Definitions Facility and Equipment Environmental Loss of facility, launch systems, and associated hardware Irreversible severe environmental damage that violates laws and regulations 2 - Critical Severe injury Major damage to facility, launch systems and associated hardware 3 - Marginal Minor injury Minor damage to facility, launch systems and associated hardware Reversible environmental damage causing a violation of law or regulation Minor environmental damage without violation of law or regulation where restoration is possible 4 - Negligible Minimal first aid required Minimal damage to facility, launch systems and associated hardware Minimal environmental damage without violating laws or regulations 99

101 Table 7.3-3: Risk Assessment Codes (RAC) Likelihood Catastrophic 2 Critical 3 Marginal A - Frequent A 2A 3A 4A B - Probable B 2B 3B 4B C - Occasional C 2C 3C 4C D - Remote D 2D 3D 4D E - Improbable E 2E 3E 4E 4 Negligible Table 7.3-4: Risk Levels Assessment Risk Levels Assessment Risk Levels Risk Assessments High Risk Highly undesirable, will lead to failure to complete the project Moderate Risk Undesirable, could lead to failure of project and loss of a severe amount of competition points Low Risk Acceptable, won t lead to failure of project but will result in a reduction of competition points Minimal Risk Acceptable, won t lead to failure of project and will result in only the loss of a negligible amount of competition points 00

102 Table 7.3-5: Launch Pad Risk Assessment Launch Pad Risk Assessment Hazard Cause Effect Pre-Mitigation RAC Failure of launch vehicle to meet stability velocity before leaving launch rail Misalignment in launch rail causing guidance pins to break or get stuck Instability of launch vehicle during launch Pre-Risk Mitigation Post-Mitigation RAC 3D Low Use correct launch lugs and firmly embed them into rocket body 3E Post- Risk Low Unstable launch pad Un-level ground Launch vehicle may leave launch platform in an unpredictable manner Launch vehicle may not reach the set competition altitude 2D Moderate Prior to launch, the launch platform should be checked for stability and correct alignment. All members present at launch should follow NAR/TRA Minimum Distance regulations 3E Low 0

103 Launch Vehicle and Recovery System Risk Assessment Table 7.3-6: Launch Vehicle and Recovery System Risk Assessment Hazard Cause Effect Pre Mitigation RAC Drogue or main parachute fails to deploy Launch vehicle is unstable after leaving launch pad Black powder charges fail to ignite Malfunction in the e-matches Malfunction in altimeters Altimeters fail to send signals Incorrect wiring of avionics and pyrotechnics Doesn t reach high enough velocity after leaving launch pad Launch vehicle motor does not have enough thrust Launch vehicle is too heavy Too much friction between launch Irreparable damage to launch vehicle, its components, and electronics Failure to meet reusability requirement Failure to meet landing kinetic energy requirement Unpredictable trajectory that could lead to crash Failure to meet altitude requirements Non-ideal launch vehicle position for drogue and main parachute deployment Pre - Risk Mitigation Post Mitigation B High Redundant black powder charges, altimeters, and e- matches Ground testing of electric ignition system (igniting black powder charges) Detailed launch procedure check list, that includes all the procedures of properly installing all avionics and pyrotechnics in the launch vehicle, will be created and followed 3E Low Create model to determine the launch vehicle s stability velocity based on fin and launch vehicle size Create model to predict launch vehicle s launch pad exit velocity and use model to select approximate motor size Perform test to determine the friction coefficient between the 2E 4D Post - Risk Low Minimal 02

104 Structural failure/shearing of fins during launch Failure of launch vehicle s internal bulkheads rails and launch vehicle Insufficient epoxy used during installation of fins Epoxy used to install fins is improperly cured Launch force on bulkheads is larger than they can support Bulkheads are poorly manufactured Unstable launch vehicle, resulting in an unpredictable trajectory Possible launch vehicle crash and injury to personnel Main and drogue parachutes attached to bulkhead will become useless Internal components supported by bulkheads will become insecure and could be damaged Damage to critical avionics systems Failure of recovery system and loss of launch vehicle launch rail and launch vehicle Use lubricant to reduce launch rail friction D Moderate Reinforce fins with sheets of carbon fiber Examine epoxy for any cracks prior to launch Perform test on fin installation Ensure all personnel are alert and are the appropriate distance away from launch pad during launch D Moderate Create prediction models of the force the bulkheads will receive during launch Use model to ensure all bulkheads are within a margin of safety Perform static load test on all bulkheads Perform detailed inspection of all manufactured bulkheads prior to launch 2E 2E Low Low 03

105 Launch vehicle motor fails to ignite Buckling of the launch vehicle s main body tube during launch Poorly installed e- match Malfunction in e- match Defective motor Body tube receives greater forces than it can support Launch vehicle will not launch Failure to meet launch requirements Structural failure of launch vehicle during flight Failure to meet launch vehicle requirements 2E Low Follow NAR safety guide lines, waiting a minimum of 60 seconds before approaching launch vehicle Once the RSO gives the all clear, check the ignition system for any loss of connection or faulty ignitors and fix connection or ignitors If problem continues, replace motor with spare E Low Create prediction models of the force the body tube will receive during launch Ensure body tube was correctly manufactured with good structural properties (correct vacuum bagging process was used in the creation of the carbon fiber) Perform static load test on the body tube 4E D Minimal Minimal 04

106 7.4 Personnel Hazard Analysis 7.4. Hazardous Materials Handling and Operations The team must be aware of and abide by these safety parameters while interacting with hazardous materials. Only designated personnel may handle these materials in accordance to state, federal, NAR, and TRA regulations Hazard Recognition and Accident Avoidance It is the safety officer s responsibility to thoroughly educate the team members with the necessary precautions when acting on behalf of the UMBRA NSL Team. The safety officer must review the safety briefings to all members before critical events such as launches and manufacturing processes to ensure the safe practice at all times. These briefings will instruct the members about the proper handling and good practices to keep and maintain the collective safety of the team. When necessary, the safety officer will only utilize experienced team members to demonstrate these practices. The safety officer will provide an electronic copy available online and upon request, physical copies of the MSDS throughout the aerospace labs. When it is deemed necessary, the safety officer has the authority to interrupt ongoing activities and lead impromptu safety mitigations. In the absence of the safety officer, participating members must review the provided MSDS and immediately notify the safety officer for documentations to prevent any further incidents. Safety offenders who violate and continue to violate the safety protocols will face disciplinary actions Carbon Fiber Composites The utilization of pre-impregnated carbon fiber for a strong body tube presents hazards to the manufacturers. Prolonged exposure to carbon fiber may lead to skin and respiratory irritation and lung disease. In addition, contact with epoxy resins can cause inflammation and conjunctivitis. Before the manufacturing process, manufacturers must read and understand the material safety data sheets for safety guidelines and mitigations. All members must wear protective gloves, safety goggles, respirators, and aprons while handling composite materials. The safety officer must be present to oversee the safety process and be prepared to handle emergency situations Other Construction Materials The operation guidelines for any construction material follows the same safety precautions as aforementioned for handling carbon fiber composites. Team members must be aware of the 05

107 consequences of handling the specified material. If the members are unaware or uncertain of the effects, they must review and understand the MSDS or communicate with the safety officer. The operating members must not handle the material in question without proper understanding of it Pre-launch Briefing All members who attend launches must attend to each prior pre-launch briefings. The safety officer will review all applicable rules and regulation in accordance to the launch site, FAA, NSL Safety Regulations, NAR and TRA Code for High Power Rocketry, and federal, state, and municipal legislation. These briefings will be unique to the specific launch and launch site and individuals who choose not to attend will be banned from the event Launches A checklist must be established and reviewed by the team before the launch event. This checklist will encompass all required materials and procedures to ensure safety while assembling and launching the launch vehicle. Before the event, the team will verify that the weather is suitable for launch and will be postponed if the weather is deemed otherwise. The safety officer must inspect and educate the process, in which designated members handle assembly, activations*-, ejection charges, and motors, prior to the installation. The launch vehicle must be assembled at their respective workplace away from the launch rails and other possible hazards. Only after completing the assemblage may the team deliver the launch vehicle to the proper launch site. Fine tuning and avionic triggers are only allowed at the launch rail when the motor and ignition system are both unarmed. The arming sequence must be performed only after all other adjustments have been completed. When the preparations are finished, the igniters may be inserted into the motor and secured into place and subsequently the ignition system must be on standby. The safety officer will ensure that all personnel are accounted for and under blockhouses or viewing bunkers. The launching sequence will begin only after the launch site is clear. If a misfire occurs, the ignition system must be disabled and a mandatory wait time must be obeyed until the safety officer determines that it is safe to approach the launch vehicle Unmanned Rocket Launches and Motor Handling Compliance with Federal, State, and Local Laws Each UMBRA NSL member must review and acknowledge the following federal, state, and local laws and regulations pertaining to amateur rocketry. Federal Aviation Regulations 4 CFR, Subchapter F, Part 0, Subpart C; Amateur Rockets: Code of Federal Regulation 27 Part 5: Commerce in Explosives; and Fire Prevention 06

108 NFPA 27 Code for High Power Rocket Motors. The team reviews and acknowledges the rules of and submits the necessary documentations for each launch site such as FAA waivers. The team s launch vehicle will adhere to the maximum altitude granted at the sites and fly only in clear and calm weather conditions. Launches will take place at:. Friends of Amateur Rocketry 2. Mojave Desert Advanced Rocketry Society 3. Lucerne Dry Lakebed Written Compliance to Safety Regulations The team is required to individually review and acknowledge the following NASA safety regulations. These regulations are also incorporated into the prerequisite team contract and subsequent activities.. A complete safety inspection, detailed in Pre-launch Checklist, will be reviewed before every launch and the team must abide by the determination of the inspection. 2. The range safety officer has authority to override the safety inspection and the launch if the launch vehicle deviates from the set guidelines and is deemed unsafe Caution Statement Plans, Procedures, and Other Working Documents Each member engaging in activities that involving materials with potential hazards are required to attend all necessary safety meetings prior to handling the materials. Any time the member is uncertain of the proper practice, the member must review the corresponding MSDS and ask for the safety officer s assistance. In addition to these safety protocols, all lab safety rules must be incorporated into the activity at all time. The safety procedures are provided below.. Gain permission from the safety officer for usage of hazardous materials a. Safety officer must approve in writing for each member 2. Verify proper attire in the lab environment. Long hair must be tied to prevent vision obstruction and tangling risks a. Closed-toed shoes must be worn at all times b. Clothing must not be too loose or constricting c. Protective gear must be worn when handling hazardous materials as instructed by the safety officer and corresponding MSDS 3. Confirm the alertness of all participating members 4. Prepare and keep the work area clean and free of obstruction 5. Obtain and handle the hazardous materials 6. Properly dispose of the excess hazardous materials and clean and organize the lab space 07

109 Table : Personal Protective Equipment Product Chemical Family Manufacturer Hazards PPE West System 05 Epoxy Resin 2 West System 205 Fast Hardener 3 Aluminosilicate Fiber Blankets (TaoFibre Blanket) 4 Dan Tack 2028 Contact Spray Super Adhesive 5 West System 05 Epoxy Resin Epoxy Resin Amine Ceramic Fiber (RCF) Aerosol Adhesive Epoxy Resin West System Inc. West System Inc. InterSource USA Inc. Adhesive Solutions Inc. West System Inc. May cause skin irritation, eye irritation, and allergic reaction. Burns to eyes and skin; harmful if swallowed or ingested. Prolonged exposure to dust may cause skin, eye, and respiratory tract irritation. May cause headaches, dizziness, unconsciousness, injury, and toxicity, skin and eye irritation. May cause skin irritation, eye irritation, and allergic reaction. Gloves, loose clothing, goggles, no exposed areas. Gloves, loose clothing, goggles, no exposed areas, face gear. Gloves, loose clothing, goggles, no exposed areas, protective breathing masks. Proper respiratory equipment and other facial gear including goggles. Gloves, loose clothing, goggles, no exposed areas. 08

110 6 Generic Oriented Strand Board Material 7 MTM49L Epoxy Resin 8 R-Matte Plus-3 (Sheathing Insulation Board) 9 West System 05 Epoxy Resin 0 Aeropoxy PH3630 Aeropoxy PH6228A N/A Epoxy Resin Polyisocyanurate Foam Epoxy Resin Modified Amine Mixture Epoxy Resin Based Mixture N/A (MSDA provided by Structural Board Association) Advanced Composites Group Inc. Rmax Operating, LLC. West System Inc. Aeropoxy Aeropoxy Inhalation and exposure to dust can cause dizziness, skin and eye irritation, serious injury, or even death. Inhalation and exposure can cause respiratory defects and skin/eye irritation, or allergic reaction. May cause skin irritation, eye irritation, and allergic reaction; known carcinogenic material (harmful in overexposure). May cause skin irritation, eye irritation, and allergic reaction May cause skin irritation, eye irritation, and allergic reaction. May cause skin irritation, eye Ventilation, Protective Gloves, Respiratory Protection Gloves, loose clothing, goggles, no exposed areas. Gloves, loose clothing, goggles, no exposed areas, proper ventilation. Gloves, loose clothing, goggles, no exposed areas. Gloves, loose clothing, goggles, no exposed areas. Gloves, loose 09

111 irritation, and allergic reaction. clothing, goggles, no exposed areas. 2 Aeropoxy PH6228B Modified Amine Mixture Aeropoxy May cause skin irritation, eye irritation, and allergic reaction. Gloves, loose clothing, goggles, no exposed areas. 3 Aeropoxy PH3660 Epoxy Resin Based Mixture, Diphenylolpropane Aeropoxy May cause skin irritation, eye irritation, and allergic reaction; liver, kidney irritation with overexposure. Gloves, loose clothing, goggles, no exposed areas, proper ventilation. 4 Aeropoxy PH3665 Modified Amine Mixture Aeropoxy Skin, Eye, and Lung irritation with overexposure; toxicity. Gloves, loose clothing, goggles, no exposed areas, proper ventilation. 5 Aeropoxy PR2032 Multifunctional acrylate Aeropoxy Skin, Eye, and Lung irritation with overexposure; toxicity. Gloves, loose clothing, goggles, no exposed areas, proper ventilation. 0

112 7.4.9 Machining Lab Machine operators on the team must be adhere to the regulations such as the yellow and red tag set forth by California State Polytechnic University. To receive the certifications, members must complete trainings administered by Professor Stover, the Director of the Engineering Project Development Laboratory, which includes a lab tour, machine demonstrations, and subsequent exam. Yellow tag allows the team to utilize drill press, saws, belt sanders, grinders, and metal brakes while red tag certifies for lathes and CNC utilization. Regardless of certifications, all operators must wear protective gear and avoid wear loose clothing and hair. Operator must be educated about the locations of safety stations throughout the lab and against working alone. Table : Lab and Machine Shop Risk Assessments Lab and Machine Shop Risk Assessment Hazard Cause Effect Pre Mitigation RAC Pre - Risk Mitigation Post Mitigation Post - Risk Personnel injury when working with chemicals Chemical spill/splash Exposure to chemical fumes Skin, eye, and lung irritation Mild to serve skin burns Lung damage or asthma 3C Moderate MSDS will be readily available in all labs at all times. They will be reviewed prior to working with any chemicals Gloves and safety glasses will be worn when handling hazardous chemicals All personnel will be familiar with locations of safety equipment including chemical showers and eye wash stations 4D Minimal

113 Personnel injury when using power tools and hand tools such as hammers, saws, and drills Personnel injury and improperly manufactured components when using lab machines such as mills, sanders, or table saws Personnel injury during carbon fiber Improper training in tool and lab equipment Lack of training in lab machine use Excess exposure to airborne Mild to severe cuts or bruises Damage to tools and equipment Damage to launch vehicle and AGSE components Mild to severe cuts and bruises Poorly fabricated parts Mild to severe irritation of 2C Moderate All personnel must be properly trained in tool use All personnel must wear safety glasses, gloves and other PPE when using tools Tools should be properly stored and taken care of Appropriate apparel must be worn when working in lab 2C Moderate All personnel working in lab must receive Yellow Tag certification before using any lab machines Personnel using more advance machines must have Red Tag certification Testing and validation of all manufactured parts must be done 3D Low Manufacture of carbon fiber must be done outside or in a well-ventilated lab 3D 3D 4E Low Low Minimal 2

114 fabrication and cutting Poorly manufactured carbon fiber components fiber particles Mishandling of epoxy and resins used Improper storage of prepreg carbon fiber leading to break down of chemical properties Leaks during vacuum bagging process Poor selection in material for the breather, release film, and sealant tape used during the vacuum skin, eyes and lungs Voids, wrinkles, and imperfections in carbon fiber Structural failure in the carbon fiber body tube Rough fin and body surfaces Misalignment of different rocket sections Proper PPE must be worn at all times including safety mask, goggles, and gloves All vacuum bagging safety procedures must be followed 3C Moderate Vacuum debulking should be performed to eliminate wrinkles and voids Vacuum requirements must be met prior to heating during the construction of carbon fiber Leak checks must be performed prior to cure and heat-up Testing and validation of all constructed carbon fiber parts must be done 3D Low 3

115 bagging process 4

116 7.5 Environmental Concerns and Effects The greatest possible threat to the well-being of the environment originates from the motor. The APCP combustion releases hazardous byproducts such as hydrochloric acid. Hydrochloric acid affects the launch rail s surroundings by changing ph of any nearby water leading to deaths in both plants and fish species. In addition, hydrochloric acid can bond with ozone producing halocarbons. These substances do not easily degrade in natural environments and as a result, tend to accumulate leading to fire hazards. However, the threats, minimal in relation to the Space Shuttle Solid Boosters, are only present in the immediate vicinity and the overall effect on the global environment is nominal. It is also important to note that testing facilities that the team will be utilizing are locations in deserted areas. There are mainly dirt and shrubs for miles around the test zones. Any emergency water supplies are safely installed in huts and at strategic locations far from the launch pad. In addition, the components of the launch vehicle are also potential concerns. In the event of parachute failure, the carbon body tube will be scattered across the area. These carbon fiber pieces, like the nylon parachutes, will not degrade and pose choking threats to wildlife. Acrylonitrile Butadiene Styrene, commonly known as ABS, is another potential area of concern. Printing with ABS, especially when overheating, may cause an accumulation of Hydrogen Cyanide (HCN). High concentration of HCN leads to various health conditions such as asthma and bronchitis. There is also a fire hazard during the process. Curing agents can further disrupt the environment. When the epoxy and hardener fuse, the mixture hardens within a short amount of time. This may disrupt and harm the environment when it is inappropriately disposed of or spilled. 5

117 8.0 Project Plan 8. Requirements Compliance Matrix The requirements compliance matrix for each subsection will be outlined in this section. The requirements compliance matrix follows the nomenclature in Table 8.-. Method of Verification Table Definition Table 8.- Method of Verification Definition Test V = Verified 2 Analysis IP = In Progress 3 Demonstration NV = Not Verified 4 Inspection 6

118 8.. Launch Vehicle Compliance Matrix Vehicle Requirements (VR) Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV The vehicle shall deliver the science or engineering payload to an apogee altitude of 5,280 feet above AGL VR. Combination of the aerodynamics of a tapered Trapizoidal fins and parabolic nose cone, as well as the thrust of an L-class motor selection and the capabilities of the structure; Launch tests shall determine whether the team shall improve on the design to ensure negligible error x x OpenRocket simulations of final design provide projected altitude, launch tests shall showcase altitude reached VR.2 VR.2. VR.2.2 VR.2.3 VR.2.4 The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in determining the award winner. Avionics subsystem will provide a Primary Perfect Flight StratoLogger The official scoring altimeter shall report the official competition Avionics subsystem will provide a altitude via a series of beeps to Primary Perfect Flight be checked after the competition StratoLogger flight. Teams may have additional altimeters to control vehicle electronics and payload experiment(s). At the LRR, a NASA official will mark the altimeter that will be used for the official scoring. At the launch field, a NASA official will obtain the altitude by listening to the audible beeps reported by the official competition, marked altimeter. Avionics Subsystem Aim USB Altimeter and RRC3 altimeters are also used Team will be on site to provide the altimeter to the NASA official After touch-down of rocket, the team will be available to provide the NASA official with the necessary altimeter x x The team shall review the launch vehicle system, subsystems, and components design verifying atleast one commercially available altimeter This will be determined during launch day, inspection of the beeps x N/A 5. x N/A x Can only be verified during flight day by NASA Official 7

119 VR.2.5 VR.2.6 VR.2.6. VR VR VR VR.3 VR.4 At the launch field, to aid in determination of the vehicle s apogee, all audible electronics, except for the official altitudedetermining altimeter shall be capable of being turned off. The following circumstances will warrant a score of zero for the altitude portion of the completion: The official, marked altimeter is damaged and/or does not report and altitude via a series of beeps after the team's competition flight. The team does not report to the NASA official designated to record the altitude with their official, marked altimeter on the day of the launch. All electronics will be capable of being shut down by discontecting batteries Purchase a durable altimeter to withstand damages Team Lead to report to the NASA Official x 5. x The altimeter reports an apogee altitude over 5,600 feet AGL Design to reach below 5,600 feet x The rocket is not flown at the competition launch site. All recovery electronics shall be powered by commercially available batteries. The launch vehicle shall be designed to be recoverable and reusable. Proper directions to the flight site arriving there on time Recovery electronics will use.v LIPO Batteries Main parachutes will allow the launch vehicle to become recoverable and all structures and electronics will be capable of being reusable The avionics team shall demonstrate during launch tests that other audible electronics have an off switch and all inaudible electronics shall not create random disturbances Inspection of the altimeter for any damage at the competition 5. Team Lead to 5. x x x 4.4. x Inspection of the altimeter for the altitude during the competition Demonstrate launch at the competition The team will inspect the recovery bay after manufacturing to ensure it is powered by batteries Conducting launch tests shall be able to result in a usable vehicle afterwards. The structures team shall design the structure to be robust and withstand impact loads through testing 8

120 VR.5 VR.6 VR.7 VR.8 VR.9 VR.0 The launch vehicle shall have a maximum of four (4) independent sections. An independent section is defined as There will be two sections in a section that is either tethered total; Module and Module 2 to the main vehicle or is recovered separately from the main vehicle using its own parachute. The launch vehicle shall be limited to a single stage. The launch vehicle shall be capable of being prepared for flight at the launch site within 4 hours, from the time the Federal Aviation Administration flight waiver opens. Only one stage will be implemented with an L-class motor located in the motor bay Launch vehicle will be fully assembled and structurally sound before hand, preflight tests and saftey tests will take minimal time The launch vehicle shall be On board electronics will be capable of remaining in launchready configuration at the pad for powered by comercially available battereies which will alow the a minimum of hour without rocket to attain functionallity for losing the functionality of any hours critical on-board component. The launch vehicle shall be capable of being launched by a standard 2 volt direct current firing system. The firing system will be provided by the NASAdesignated Range Services Provider. The launch vehicle shall use the Aerotech L50 rocket motor The launch vehicle shall require no external circuitry or special The only firing system used will ground support equipment to be the one provided by NASA initiate launch (other than what is Range Services Provider provided by Range Services) x 4.3. x 5. x 5. x 5. x 5. x The team shall review the system design making sure there are less than four independent sections During the design phase, the team lead shall inspect the design to be only single stage Before test launches, the team shall time the preperation of the launch vehicle Before test launches, the team shall demonstrate if the vehicle can be ready on the stand for hour Launch tests shall be done with a 2 V firing system proving its ability to function Inspection of the electrical circuitry design 9

121 VR. VR.2 The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR). Pressure vessels on the vehicle shall be approved by the RSO and shall meet the following criteria: The launch vehicle shall use the Aerotech L50 rocket motor 4.3. x Inspection of the purchased motor to be correct VR.2. The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) shall be 4: with supporting design documentation included in all milestone reviews. VR.2.2 VR.2.3 VR.2.4 The low-cycle fatigue life shall be a minimum of 4:. Each pressure vessel shall include a solenoid pressure relief valve that sees the full pressure of the tank. Full pedigree of the tank shall be described, including the application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when. 20

122 VR.4 VR.5 VR.6 VR.6. VR.6.2 The launch vehicle shall have a minimum static stability margin of 2.0 at the point of rail exit. The launch vehicle shall accelerate to a minimum velocity of 52 fps at rail exit. All teams shall successfully launch and recover a subscale model of their rocket prior to CDR. The subscale model should resemble and perform as similarly as possible to the fullscale model, however, the fullscale shall not be used as the subscale model. Vehicle shall have sufficient aerodynamic stablity provided by the nose cone and fins and the weight distribution to be optimized A large thrust to Weight Ratio will provide the necessary velocity A subscale model is planned for a :2 ratio; subscales launch dates are to be held weeks before CDR A subscale model is planned for a :2 ratio The subscale model shall carry an altimeter capable of reporting the Primary Perfect Flight model s apogee altitude. StratoLogger x x 5. x 4 x x Preliminary OpenRocket simulations shall show a static stability margine above 2.0 Simulations shall be conducted to analyze the minmum velocity The team shall build and test this subscale model, success criteria includes recovery system functionality and a recoverable vehicle after landing Anaylsis on performance and scaling of subscale using graphs and tables to determine comparisons Inspection that the subscale design has an altimeter VR.7 All teams shall successfully launch and recover their full-scale rocket prior to FRR in its final Gnatt Chart shows a schedual of flight configuration. The rocket when this will be acomplished flown at FRR must be the same rocket to be flown on launch day. 5. x Launch vehicle must recoverable after the launch tests conducted VR.7. The vehicle and recovery system shall have functioned as designed. TBD on flight day 5. x During flight day, the launch vehicle needs to demonstrate the objectives and operations planned 2

123 VR.7.2 VR.7.3 The payload does not have to be flown during the full-scale test flight. The following requirements still apply: () If the payload is not flown, mass During launch test, put a mass simulators shall be used to representative of the payload in simulate the payload mass. (2) the payload fairings The mass simulators shall be located in the same approximate location on the rocket as the missing payload mass. If the payload changes the external surfaces of the rocket (such as with camera housings or external probes) or manages the total energy of the vehicle, those systems shall be active during the full-scale demonstration flight. There will be minimal changes to the external sufaces of the rocket 5. x Inspection that a mass simulation is present during launch test 6 x Demonstration during test flight VR.7.4 VR.7.5 VR.7.6 The full-scale motor does not have to be flown during the fullscale test flight. The vehicle shall be flown in its fully ballasted configuration during the full-scale test flight. After successfully completing the full-scale demonstration flight, the launch vehicle or any of its components shall not be modified without the concurrence of the NASA Range Safety Officer (RSO). A smaller motor will most likely be used to save on cost, but the same type of motor might still be implemented Designed to be in fully ballasted configuration during test flight No design modification after last full-scale flight 5. x 5. x 5. x Inspection of the purchased full scale motor Will be determinded during flight day by demonstrating vehicle to be fully ballasted Inspection of rocket after full scale demonstration shall have the same results as before flight day. 22

124 VR.7.7 VR.8 VR.9 VR.9. VR.9.2 VR.9.3 VR.9.4 VR.9.5 VR.9.6 VR.9.7 Full scale flights must be completed by the start of FRRs (March 6th, 207). If the Student Launch office determines that a Gnatt Chart shows a schedule of re-flight is necessary, than an when this will be acomplished extension to March 24th, 207 will be granted. This extension is only valid for re-flights; not first time flights. Any structural protuberance on the rocket shall be located aft of the burnout center of gravity. Vehicle Prohibitions The launch vehicle shall not utilize forward canards. There will be no changes to the external sufaces of the rocket x Inspection of the Gantt Chart x There are none 4.2. x The launch vehicle shall not utilize forward firing motors. There are none 4.3. x The launch vehicle shall not utilize motors that expel titanium Aerotech L50 rocket motor does sponges (Sparky, Skidmark, not utilize those features 4.3. x MetalStorm, etc.) The launch vehicle shall not Aerotech L50 rocket motor is a utilize hybrid motors. solid motor 4.3. x The launch vehicle shall not utilize a cluster of motors. Only one motor will be used 4.3. x The launch vehicle shall not utilize friction fitting for motors. Bulkheads and motor casing will be securly fastening into place and not friction fit The launch vehicle shall not Weight of vehicle shall counter exceed Mach at any point during balance with the thrust to ensure flight. excessive amounts of power to enable speed of mach x x Inspection of the structural design and pertubations Inspection of the structural design and pertubations Inspection of the propulsion system and the final rocket configuration Team shall investigate and research on the details of the motor Team shall demonstrate the use of a solid motor during tests Inspection of the propulsion system and the final rocket configuration During manufacturing the team shall not use friction fitting and demonstrate the vehicle can function with other means From OpenRocket simulations and constant iterations of designs, vehicle design to never reach Mach 23

125 VR.9.8 notes: VR.3 is removed Vehicle ballast shall not exceed 0% of the total weight of the rocket Recovery System Compliance Matrix TBD on flight day 4.2. x x The structures team shall inspect the vehicle and confirm that the vehicle ballast is less than 0% of the total weight Totals Recovery System Requirements Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV RSR2. The launch vehicle shall stage the deployment of its recovery devices, Piston ejection system for main parachute and black powder ignition for drogue Recovery testing will be done to determine the proper where a drogue parachute is deployed at deployment deployment of parachutes apogee and a main parachute is deployed 4.4. x at a much lower altitude. RSR2.2 RSR2.3 RSR2.4 RSR2.5 RSR2.6 Each team must perform a successful ground ejection test for both the drogue and main parachutes. This must be done prior to the initial subscale and full scale launches. At landing, each independent sections of the launch vehicle shall have a maximum kinetic energy of 75 ft-lbf. The recovery system electrical circuits shall be completely independent of any payload electrical circuits. Recovery tests and ejection tests done on 2/3/206 for subscale and 2/4/207 for full scale 8.3. x Custom drogue parachute and custom main parachute Recovery bay will hows recovery system electrical circuits in between main and drogue bay while payload electrical circuits will be housed in bayload bay below the drogue bay The recovery system shall contain Perfectflite stratologger AltimeterCF will be redundant, commercially available used to measure the rockets altitude and set altimeters. The term altimeters includes off the deploy charges at programmed both simple altimeters and more altitude sophisticated flight computers. Motor ejection is not a permissible form of primary or secondary deployment x x x Motor ejection will not be part of the recovery design 4.4. x x Results of the ground ejection test shall verify successful performance Hand calculations and respective simulations to analyze kinetic energy The launch vehicle shall demonstrate recovery system is separate from electrical circuits during launch test Inspection of the altimeters chosen in the design Inspection of vehicle design, and demonstration of launch test 24

126 RSR2.7 RSR2.8 RSR2.9 RSR2.0 Each altimeter shall be armed by a dedicated arming switch that is accessible from the exterior of the rocket airframe when the rocket is in the launch configuration on the launch pad. Each altimeter shall have a dedicated power supply. Each arming switch shall be capable of being locked in the ON position for launch. Removable shear pins shall be used for both the main parachute compartment and the drogue parachute compartment. RSR2. An electronic tracking device shall be installed in the launch vehicle and shall transmit the position of the tethered vehicle or any independent section to a ground receiver. RSR2.. Any rocket section, or payload component, which lands untethered to the launch vehicle, shall also carry an active electronic tracking device. There will be hatches to access the altimeters in the recovery bay for both subscale and full scale.v Lipo Batteries will provide power for each altimeter Design a rotary switch Parachute Bay design shall have shear pins TELEGPS will be used to send real time gps data to ensure rocket recovery and the TELEGPS Starter pack with dongle will plug into a laptop to receive data All sections, including payload components are tethered to launch vehicle x x x x x x RSR2..2 The electronic tracking device shall be Tests will be done with the GPS prior to fully functional during the official flight on launch day x launch day. RSR2.2 The recovery system electronics shall not be adversely affected by any other onboard electronic devices during flight (from launch until landing). Recovery system electronics shall have independent components x x Inspection of the design of the rocket inspection of altimeter and avionics schematic Demonstration of the arming switch during manufacturing Inspection of the parachute bay design, and construction Inspection of th avionics schematic Inspcetion of design to have no untethered sections Demonstration that tracking device works prior to launch day Demonstration that all recovery systems electronics are independent during recovery test 25

127 RSR2.2. The recovery system altimeters shall be physically located in a separate compartment within the vehicle from any other radio frequency transmitting device and/or magnetic wave producing device Payload Compliance Matrix Recovery system altimeters shall be in the recover bay, separate from other devices that can interfere with it RSR2.2.2 The recovery system electronics shall be Recovery system bay will provide protection shielded from all onboard transmitting devices, to avoid inadvertent excitation of the recovery system electronics. RSR2.2.3 The recovery system electronics shall be Recovery system bay will provide protection shielded from all onboard devices which may generate magnetic waves (such as generators, solenoid valves, and Tesla coils) to avoid inadvertent excitation of the recovery system. RSR2.2.4 The recovery system electronics shall be Recovery system bay will provide protection shielded from any other onboard devices which may adversely affect the proper operation of the recovery system electronics. notes x x x x Inspection of the recovery bay design Inspection of recovery bay design Inspection of recovery bay design Inspection of recovery bay design Totals 2 5 Experiment Requirements Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV Each team shall choose one design experiment 3.3 and 3.4 Investigation of the experiment ER3.. experiment option from the following list. 6. dcumentation in PDR x Additional experiments (limit of ) are experiment 3.4 Investigation of the experiment ER3..2 encouraged, and may be flown, but they will 6.2 chosen in PDR not contribute to scoring. x ER3..3 If the team chooses to fly additional Results of additional experiments in experiments, they shall provide the PDR and CDR and LRR appropriate documentation in all design reports so experiments may be reviewed for flight safety. 6.2 x Investigation of the experiment chosen in PDR 26

128 Experiment Requirements Option 2 Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV ER3.3 Roll induction and counter roll Teams shall design a system capable of RIS - A (Inertia Flywheel Design), RIS - Inspection of payload operations ER3.3. controlling launch vehicle roll post motor B (Fin Aileron Design), or RIS - C 6..,6..2 and design to function post burnout. (Aerofan Design) will begin at post burnout x x ER3.3.. ER ER ER3.3.2 ER3.3.3 The systems shall first induce at least two rotations around the roll axis of the launch vehicle. After the system has induced two rotations, it must induce a counter rolling moment to halt all rolling motion for the remainder of launch vehicle ascent. Teams shall provide proof of controlled roll and successful counter roll. Teams shall not intentionally design a launch vehicle with a fixed geometry that can create a passive roll effect. Teams shall only use mechanical devices for rolling procedures. motor burnout RIS - A (Inertia Flywheel Design) uses moment of inertia of heavy cylindrical object, RIS - B (Fin Aileron Design) uses aerodynamics manipulation to roll, or RIS - C (Aerofan Design) uses aerodynamic manipulation RIS - A (Inertia Flywheel Design) uses moment of inertia of heavy cylindrical object, RIS - B (Fin Aileron Design) uses aerodynamics manipulation to roll, or RIS - C (Aerofan Design) uses aerodynamic manipulation Redundant data retrieval: live acceleromater data for quantitative represenation, recorded footage post launch RIS - A (Inertia Flywheel Design), RIS - B (Fin Aileron Design), and RIS - C (Aerofan Design) are all passive effects RIS - A (Inertia Flywheel Design), RIS - B (Fin Aileron Design), and RIS - C (Aerofan Design) are all using mechanical devices x x x x x Testing of rolling moment Testing of rolling moment Demonstration of capabilities of data retrieval and cameras for footage Inspection of chosen payload design Inspection of chosen payload design 27

129 Experiment Requirements Option 3 Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV ER3.4 Fragile material protection Teams shall design a container capable of Alternative, 2, and 3 designs Different materials placed inside ER3.4. protecting an object of an unknown material accommodate for unknown sizes and payload and determine if it and of unknown size and shape. shapes in the container survives a drop test x ER3.4.. ER ER ER There may be multiple of the object, but all copies shall be exact replicas. The object(s) shall survive throughout the entirety of the flight. Teams shall be given the object(s) at the team check in table on launch day. Teams may not add supplemental material to the protection system after receiving the object(s). Once the object(s) have been provided, they must be sealed within their container until after launch. Container shall be able to hold multiple object safety, in a compartment or case The usage of a soft material will allow for load absorption and the encasing device will secure the object in place The design shall accommodate all shapes and sizes and strengths of the material Design shall have an encasing device for objects and no removable parts x x x x Different amount of materials placed inside payload and determine if it survives a drop test Materials placed inside payload shall survive during drop tests Demonstration that the payload may accommodate any object Demonstration that the payload shall not need any removable mechanisms or any removable items within its structure to perform ER ER notes The provided object can be any size and shape, but will be able to fit inside an imaginary cylinder 3.5 in diameter, and 6 in height. The object(s) shall have a maximum combined weight of approximately 4 Dimensions shall be larger than 3.5" in diameter and 6" in heightto accommodate the material Container shall be able to withstand 4 ounces x x x Inspection of design to fit the dimensions listed earlier Testing of payload performance for materials wieghing 4 ounces inside Totals

130 8..4 Safety Compliance Matrix Safety Requirements Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV Each team shall use a launch and safety checklist. The final checklists shall be Safety Managers and Officers will create a checklist prior to FRR and LRR At LRR, demonstration of the use of the checklist SR4. included in the FRR report and used during the Launch Readiness Review (LRR) and any 7. launch day operations. x SR4.2 Each team must identify a student safety Michael Nguyen Demonstrated in team 7. officer who shall be responsible for all items x description of PDR SR4.3 The role and responsibilities of each safety officer shall include, but not limited to: SR4.3. Monitor team activities with an emphasis on SR4.3.. Design of vehicle and launcher Outlined for Safety Officer responsibilities Demonstrate safety during 7. x design SR Construction of vehicle and launcher Outlined for Safety Officer responsibilities Demonstrate safety during 7. x construction SR Assembly of vehicle and launcher Outlined for Safety Officer responsibilities Demonstrate safety during 7. x assembly SR Ground testing of vehicle and launcher Outlined for Safety Officer responsibilities Demonstrate safety during 7. x ground testing SR Sub-scale launch test(s) Outlined for Safety Officer responsibilities Demonstrate safety during subscale launch test 7. x SR Full-scale launch test(s) Outlined for Safety Officer responsibilities Demonstrate safety during fullscale launch test 7. x SR Launch day Outlined for Safety Officer responsibilities Demonstrate safety during 7. x launch day SR Recovery activities Outlined for Safety Officer responsibilities Demonstrate safety during 7. x recovery test 29

131 SR SR4.3.2 SR4.3.3 Educational Engagement Activities Implement procedures developed by the team for construction, assembly, launch, and recovery activities Manage and maintain current revisions of the team s hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data Outlined for Safety Officer responsibilities 7. Safety Manager and Team Lead will lead the team to follow all procedures made 7. Safety Officer will update the hazard, failure, procedure, and MSDS sheets for all reviews in accordinance to new materials and regulations 7. x x Demonstrate safety during educational engagement Safety task for to demonstrate these procedures Review (preliminary, critical, etc.) documents will demonstrate these hazard analyses, failure modes, and procedures SR4.3.4 Assist in the writing and development of the team s hazard analyses, failure modes analyses, and procedures. Safety task force are in charge of writing proposal section 3.0; safety plans, procedures, and other documents 7. x Review (preliminary, critical, etc.) documents will demonstrate these hazard analyses, failure modes, and procedures SR4.4 Each team shall identify a mentor. A Rick Maschek mentor is defined as an adult who is included as a team member, who will be supporting the team (or multiple teams) throughout the project year, and may or may not be affiliated with the school, institution, or organization. The mentor shall maintain a current certification, and be in good standing, through the National Association of Rocketry (NAR) or Tripoli Rocketry Association (TRA) for the motor impulse of the launch vehicle, and the rocketeer shall have flown and successfully recovered (using electronic, staged recovery) a minimum of 2 flights in this or a higher impulse class, prior to PDR. The mentor is. x x Mentor demonstrates agreement to work with the team 30

132 SR4.5 SR4.6 notes During test flights, teams shall abide by the rules and guidance of the local rocketry club s RSO. The allowance of certain vehicle configurations and/or payloads at the NASA Student Launch Initiative does not give explicit or implicit authority for teams to fly those certain vehicle configurations and/or payloads at other club launches. Teams should communicate their intentions to the local club s President or Prefect and RSO Teams shall abide by all rules set forth by the FAA. Team members will sign a document stating that they will abide to these rules Team members will sign a document stating that they will abide to these rules x x Team members demonstrate that they signed the rules and guidance of the local rocketry club's RSO and have available by CDR Team members demonstrate that they signed the rules for FAA and have available by CDR Totals

133 8..5 General Compliance Matrix General Requirements Verification Method STATUS Design Requirements Section Verification Details REQ# Description V IP NV Students on the team shall do 00% of the project, including design, construction, written reports, presentations, and flight preparation with the Team lead will distribute equal amount of work to each student, both for writing and manufacturing Team members shall demonstrate 00% of the project GR5. exception of assembling the motors and handling black powder or any variant of ejection charges, or preparing and installing electric matches (to be done.2 x by the team s mentor). GR5.2 GR5.3 The team shall provide and maintain a project plan to Project plan to be discusses in detail in PDR, include, but not limited to the following items: project CDR, and LRR milestones, budget and community support, checklists, personnel assigned, educational engagement events, and risks and mitigations. Foreign National (FN) team members shall be CPP NSL does not have any FNs on the identified by the Preliminary Design Review (PDR) and current team may or may not have access to certain activities during launch week due to security restrictions. In addition, FN s may be separated from their team during these activities. 8.0 x.2 x Demonstration of the project plan Inspection of all team members GR5.4 The team shall identify all team members attending launch week activities by the Critical Design Review (CDR). Team members shall include: GR5.4. Students actively engaged in the project throughout the entire year. GR5.4.2 One mentor (see requirement 4.4). GR5.4.3 No more than two adult educators. All team members have their subteam and designated responsibilities.2 x Dr. Donald Edberg Dr. Todd Coburn and Rick Maschek. x. x Members shall demonstrate their engagement Demonstration of the active advisor Demonstration of the active mentors 32

134 GR5.5 GR5.6 GR5.7 GR5.8 GR5.9 GR5.0 GR5. GR5.2 GR5.3 notes The team shall engage a minimum of 200 participants in educational, hands-on science, technology, engineering, and mathematics (STEM) activities, as defined in the Educational Engagement Activity Report, by FRR. An educational engagement activity report shall be completed and submitted within two weeks after completion of an event. A sample of the educational engagement activity report can be found on page 28 of the handbook. Outreach Manager, Diran, will be in charge of planning activites with over 200 students The team shall develop and host a Web site for project cpprocketry.net documentation. Teams shall post, and make available for download, Web Manager, Shannen, will upload all the required deliverables to the team Web site by the required documents on the website by due due dates specified in the project timeline. date Documents saved in PDF format All deliverables must be in PDF format. In every report, teams shall provide a table of contents including major sections and their respective sub-sections. In every report, the team shall include the page number at the bottom of the page. The team shall provide any computer equipment necessary to perform a video teleconference with the review board. This includes, but not limited to, a computer system, video camera, speaker telephone, and a broadband Internet connection. If possible, the team shall refrain from use of cellular phones as a means of speakerphone capability. All teams will be required to use the launch pads provided by Student Launch s launch service provider. No custom pads will be permitted on the launch field. Launch services will have 8 ft. 00 rails, and 8 and 2 ft. 55 rails available for use. Teams must implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards (36 CFR Part 94) Subpart B-Technical Standards ( Software applications and operating systems Web-based intranet and Internet information Table of contents will follow cover page 9 x.4 x.4 x.4 x pg 2 Page numbers will be visible on the bottom right hand side of each review ALL x CPP NSL shall be campus during video teleconferences, with access to a computer system, projector, wifi, and other necessary communication devices N/A x Final vehicle design shall allow the use of the provided launch pads Team will implement the EIT Accessibility Standards 5. x 5. x x Demonstration of education activities Inspection of website existence Inspection of deliverables on website after each deadline Inspection of all deliverables Demonstration of tables of contents Demonstration of all page numbers Demonstration of use of computer equipment during teleconferences Demonstration during flight day to use launch pads provided Demonstration to include EIT Standards in systems Totals

135 8..6 Derived Requirements Compliance Matrix Derived Requirements (DR) Design Verification Method STATUS Section Verification Details REQ# Description Requirements V IP NV DR.0 Roll Maneuver must follow Roll Maneuver Payload tests will sequence of events: sequence of verify that the DR.0. Motor burn out events outlined sequence of events DR.0.2 On board instrumentation accounts in Review are followed for natural rotation of rocket documents and DR.0.3 The roll system shall induce a design the system 6.4 x moment to generate at least 2 full to meet the DR.0.4 After full rotation, the roll system events induces a moment to counter The system shall return the rocket to DR.0.5 its initial rotation measured at rocket burnout DR2.0 Ouline Safety Officer Responsibilities 7. x Inspection of the PDR notes Generate a list of responsibilities in PDR Totals

136 8.2 Preliminary Budget Plan 8.2. Full Scale Launch Vehicle Expenses Purpose Component Description Mass (lbs) Supplier Qty Unit Price Total Price Motor Propellant Aerotech L siriusrocketry 3 $69.99 $ Motor Casing 75/ apogee $ $ Forward Closure 75mm 0.44 apogee $0.65 $0.65 Aft Closure 75mm siriusrocketry $80.25 $80.25 Retainer 75mm Flanged apogee $53.50 $53.50 Lube Super Lube Grease 3oz 0 siriusrocketry $5.99 $5.99 Observation Bay Raspberry Pi 2 Mod. B Microcontroller for Adafruit 2 $40.00 $80.00 camera Adafruit Raspberry Pi Camera for flight Adafruit 2 $30.00 $60.00 camera footage Use to connect U-bolt 5/6" x.375" x 2.5" home depot 4 $.30 $5.20 bulkhead to parachute Used to connect modules before ejection charge Nylon Shear Pins (20 pack) Apogee Components $2.95 $2.95 Recovery bay Flange connection device PVC screws connector -/2 in x -/4in, ABS Home Depot 2 $2.48 $4.96 Provides structural support for bulkheads L-Brackets Launch Vehicle Structure Budget Stanley-National Hardware 2-Pack.5- in Metallic Corner 3D Printer Filament Black PLA.75mm kg Stanley-National via Lowes 4 $.78 $24.92 Used for the transition 3D printer filliment piece/ observation Amazon $9.95 $9.95 Support for transition Aluminum sheet 24" x 24" sheet mcmaster $33.28 $33.28 piece metal Used for transparent plexiglass sheet OPTIX 0.08-in x 8-in x lowes $2.98 $2.98 window for camera 0-in Clear Acrylic Sealing the window to silicon gel GE Iron Grip Silicone lowes $7.98 $7.98 the body tube Adhesive - Clear Body tube pre 72", 6" blue tube Blue Tube " alwaysreadyrocketry $05.95 $05.95 Connecting ring for 6" blue tube coupler Blue Tube 2.0 2" alwaysreadyrocketry 2 $9.95 $39.90 modules Length The screws used to Screws Zinc-Plated Alloy Steel Bolt Depot $.69 $.69 connect various Flat-Head Cap Screw, components #0-32 Thread, 3/4" Length (25 pack) securing devices used Nuts Hex machine screw McMaster-Carr 2 $7.78 $5.56 to hold screws in place nuts, Zinc plated steel, Epoxy used to join Epoxy 05 Resin (26.6 fl oz) West Marine $99.99 $99.99 material together Hardener 207 Hardener (27.5 fl West Marine $47.99 $47.99 oz) Used to strengthen Carbon Fiber CYCOM 5320 Epoxy Cytec $50.00 $50.00 parts and materials Resin Prepreg System Used as a structural Bulkheads /4 in x 4 ft x 8 ft Birch.323 Lowes $27.47 $27.47 support for the main Plywood body tube Total Mass.765 Pounds Total Cost $,

137 8.2.2 Sub Scale Launch Vehicle Expenses Purpose Component Description Mass (lbs) Supplier Qty Unit Price Total Price Motor Propellant Aerotech L siriusrocketry 2 $69.99 $99.99 Motor Casing w/ 54/ apogee $47.30 $47.30 forward seal disk Aft Closure 75mm siriusrocketry $63.00 $63.00 Retainer 75mm Flanged apogee $53.50 $53.50 Lube Super Lube Grease 3oz 0 siriusrocketry $5.99 $5.99 Use to connect Eye-bolt bulkhead to parachute Used to connect modules before ejection charge Recovery bay Flange connection device Provides structural support for bulkheads Used for the transition piece/ observation Support for transition piece Body tube pre transition Body tube post transition Connecting ring for modules The screws used to connect various components securing devices used to hold screws in place Used as a structural support for the main body tube Nylon Shear Pins (20 pack) Subscale Launch Vehicle Structure Budget Stanley-National Hardware 3/8-in to 8 x in Zinc-Plated Plain Eye Bolt Use from full scale pack lowes 4 $0.95 $ E-05 Apogee Components $3.0 $3.0 PVC screws connector -/2 in x -/4in, ABS Home Depot 2 $.87 $3.74 L-Brackets Stanley-National Hardware 2-Pack.5- in Metallic Corner 3D printer filliment 3D Printer Filament Black PLA.75mm kg Aluminum sheet Use from full scale metal piece 48", 3" blue tube Blue Tube " Length 48", 2.5" blue tube Blue Tube 2.0.5" Length 3" blue tube coupler Blue Tube 2.0 8" Length Screws Zinc-Plated Alloy Steel Flat-Head Cap Screw, #0-32 Thread, 3/4" Length (25 pack) Nuts Hex machine screw nuts, Zinc plated steel, Bulkheads /4 in x 4 ft x 8 ft Birch Plywood Stanley-National via Lowes 4 $.78 $ Amazon $9.95 $ mcmaster $33.28 $ alwaysreadyrocketry $29.95 $ alwaysreadyrocketry $23.95 $ alwaysreadyrocketry $9.95 $ bolt depot $.69 $ mcmaster $7.78 $ lowes $27.47 $27.47 Total Mass Pounds Total Cost $

138 8.2.3 Recovery System Expenses Purpose Component Description Mass (g) Supplier Qty Unit Price Total Price Full Scale Parachute 3000 Denier 63'' wide kevlar Ebay 2 $24.99 $2.00 Firewall Paracord Vengance Nylon wound 550 paracord paracord galaxy 2 $4.99 $9.98. Oz Nylon Rip stop Foliage Green ripstopbytheroll 25 $4.40 $0.00 Recovery. Oz Nylon Rip stop Blaze Yellow ripstopbytheroll 25 $4.40 $0.00 Parachute 5/8'' shock cord 5/8'' option, ideal for 30lbs fruitychutes 20 $.50 $30.00 System /4 inch stainless steel up to 200lb fruitychutes $5.00 $5.00 quicklink lb Rosco Swivel up to 500lb, set of fruitychutes $2.00 $2.00 TELEGPS To send real time gps data to ensure rocket recovery Apogee rockets $24.00 $24.00 Recovery Avionics Parachute Firewall Recovery Parachute System Recovery Avionics TELEGPS Starter pack with dongle Perfectflite stratologger AltimeterCF Plug into laptop to receive data transmitted by Telegps measure the rockets altitude and set off the deploy charges at V Lipo Batteries Power Source for Avionics Package Sub Scale 3000 Denier 63'' wide kevlar Apogee rockets $07.00 $07.00 Apogee rockets 2 $58.80 $7.60 Amazon 5 $20.00 $00.00 ebay 2 $24.99 $ Paracord Vengance Nylon wound 550 paracord paracord galaxy 2 $4.99 $9.98. Oz Nylon Rip stop Foliage Green ripstopbytheroll 25 $4.40 $0.00. Oz Nylon Rip stop Blaze Yellow ripstopbytheroll 25 $4.40 $0.00 5/8'' shock cord 5/8'' option, ideal for 30lbs fruitychutes 20 $.50 $30.00 /4 inch stainless steel up to 200lb fruitychutes $5.00 $5.00 quicklink lb Rosco Swivel up to 500lb, set of fruitychutes $2.00 $2.00 TELEGPS To send real time gps data Apogee rockets $24.00 $24.00 to ensure rocket recovery TELEGPS Starter pack with Plug into laptop to receive Apogee rockets $07.00 $07.00 dongle data transmitted by Telegps Perfectflite stratologger AltimeterCF Recovery System Budget measure the rockets altitude and set off the deploy charges at Apogee rockets 2 $58.80 $7.60.V Lipo Batteries Power Source for Avionics Amazon 5 $20.00 $00.00 Package Total Mass.7368 pounds Total Cost $,

139 8.2.4 Payload Expenses Purpose Component Description Mass (lbs) Supplier Qty Unit Price Total Price Surgical Tubing Strong tubing to secure FMP Amazon $22.8 $22.8 box in place White Printer Filament Filament used for printing Amazon $22.99 $22.99 FMP case Egg Crate Foam firm cushion to hold FMP in Samys $20.00 $20.00 place Plywood secure tubing and maintain Home Depot 4 $0.40 $.60 shape PERFECTFLITE USB DATA TRANSFER KIT to interface with thealtimeter and test,adjust altitude deployment, and Apogee rockets $34.6 $34.6 recover data. 0 Arduino Mega actively computes inputs and sends out signals to adafruit $45.95 $45.95 motors and antennas Adafruit 0-DOF IMU to measure the roll and Adafruit $30.00 $30.00 breakout sensor acceleration of the rocket XBee Shield Arduino - XBee interface Arduino $4.00 $4.00 (includes SD card slot) RCS Avionics Adafruit Ultimate GPS module GPS Antenna 28dB active Payload Experiment(s) Budget Recieves signal from GPS satellites and triangulates coordinates Allows the GPS Module to receive GPS signal Adafruit $30.00 $30.00 Adafruit $3.00 $3.00 XBee Pro 900 RPSMA Flight data transmitter Adafruit $40.00 $40.00.V Lipo Batteries Power Source for Avionics Amazon 5 $20.00 $00.00 Package Mhz XBee Antenna Range extender for XBee Sparkfun $7.95 $ V Lipo Batteries Power Source for Roll Amainhobbies 2 $63.00 $ Hall effect sensor Independent RPM Robotshop 4 $0.95 $3.80 measurement of RIS Small magnets RPM measurment of RIS Robotshop 4 $0.95 $3.80 RIS Motors (24V; ~0 oz-in Propell the reaction wheels - Anaheim 2 $24.00 $ torque) BLY7MDA2S-24V W automation Kapton Tape fastening and electrical Amazon $2.00 $2.00 insulator 0 Used for weight support of 5.9" OD 3.9" ID amazon 2 $0.37 $20.74 Lazy Susan bearings flywheel Steel Tube Used for the flywheel 5.5" OD 4.5" ID ' Long mcmaster $30.00 $30.00 Aluminum plate Rotate flywheel 6" x 6" Sheet cut into 3" mcmaster $7.22 $7.22 diameter cog Total Mass pounds Total Cost $,

140 8.2.5 Educational Engagement Expenses and Other Educational Engagement Budget Purpose Component Description Supplier Qty Unit Price Total Price Rocket Kit (24 Beginner rocket model Pack) kit Apogee Rockets $84.75 $84.75 Rocket Motors (3 Pack) B4-4 Motor Apogee Rockets 8 $.55 $92.40 Activity Sky Complete Launch Pad and Launch System Controller Apogee Rockets $25.94 $25.94 Eggs 8 ct Ralphs 2 $4.99 $9.98 Egg Drop Nylon Fabric white, by the yard Amazon 4 $5. $20.44 string cotton twine Amazon 2 $4.93 $9.86 Scissors kid safe dollar store 0 $.00 $0.00 Presentations trifold board white Staples $7.99 Total Cost $7.99 $36.36 Purpose Description Supplier Website Qty Unit Price Total Price Website website for CPP NSL weebly weebly $96.00 Total Cost $96.00 $ Travel Expenses Other Budget Travel Budget Purpose Description Qty Unit Price Total Price Delta Airplane Tickets 4/5/206-4/0/206 8 $ $4, Hotel Double Room 6 days, Embassy Suites and Spa, Huntsville, AL 5 $43.00 $4, Gas 4/2/206-4/0/206 Driving for 2 cars; miles $0. Total Cost $ $9,

141 8.2.7 Overall Expenses Overall Budget Launch Vehicle Structure Budget Subscale Launch Vehicle Structure Budget Recovery System Budget Payload Experiment(s) Budget Educational Engagement Budget Other Budget Travel Budget TOTAL Full Scale Launch Vehicle cost TOTAL Sub Scale Launch Vehicle cost TOTAL ALL Cost

142 8.2.8 Funding Plan The NSL CPP Team plans to acquire funds to pay for all expenses. All funding sources from previous years will be pursued once more for the project. Potential funding sources are listed below in table A great amount of funding will be through the college. Table 6.3-: Funding Sources and Support Funding Source Potential Support Cal Poly Pomona Associated Students Incorporated (ASI) Grant $5, Cal Poly Pomona Engineering Council Special Projects Funding $ Space Grant $4, Cal Poly Pomona Research and Project Grants $, Total $0,800 These funds do not provide for all the expenses; therefore, the team will reach out to local businesses and organizations for support as mentioned in 6.4. Brochures will be provided for these potential sponsors in hopes to gain their interest in the project. 8.3 Preliminary Timeline 8.3. Timeline Figure and Figure outlines the milestones main tasks throughout the competition. It is important to follow this plan as it ensures the team to be on track for a successful launch vehicle flight. Figure mentions these milestones: (WE) Website Established, (SLVC) Subscale Launch Vehicle Completion, (PC) Payload Completion, (FLVC) Full Scale Launch Vehicle Completion, (PDR) Preliminary Design Review, (CDR) Critical Design Review. (FRR) Flight Readiness Review Project Life Cycle Phases Phase A: Concept Studies and Development Phase B: Preliminary Design and Technology Completion Phase C: Final Design and Fabrication Phase D: System Assembly, Integration, Test, and Launch Phase E: Operations and Sustainment Phase F: Closeout Major Events WE SLVC PC FLVC Launch NASA Student Launch Reviews Proposal PDR CDR FRR PLAR Figure Phase Timeline 4

143 Figure Overall Timeline 42

144 8.3.2 Gantt Chart Figure Gantt Chart 43

145 44